The fuselage is the main structure or body of the aircraft and carries the passengers
and crew in safe, comfortable conditions. It also provides space for controls, accessories
and other equipment. It transfers loads to and from the main planes or wings, the tail
plane, fin, landing gear and, in certain configurations, the power plants.
There are three main types of construction in use. First there is the truss or framework
type, which is generally used for light, non-pressurised aircraft. Then there is monocoque.
And finally, semi-monocoque, which is in use on most other aircraft. The last two types
of structure are more generally referred to as stressed skin.
The framework consists of light gauge steel tubes, welded together to form a space frame
of triangular shape to give the most rigid of geometric forms.
Each tube carries a specific load, the magnitude of which depends on whether the aircraft is
airborne or on the ground. It's a strong, easily constructed and relatively trouble-free
basic structure.
The framework is covered by a lightweight aluminium alloy or fabric skin to give an
enclosed, aerodynamically efficient load-carrying compartment.
Monocoque is a French word meaning single shell. In a monocoque structure, all the loads
are taken by a stressed skin with just light internal frames or formers to give the required
shape. Sandwich construction, which is a honeycomb core with a skin of aluminium alloy or composite
material, such as glass reinforced plastic or carbon fibre reinforced plastic, can be
used to provide rigidity and strength. Even slight damage to the skin can seriously weaken
the structure.
To be a true monocoque, the structure would have no apertures in it at all, but for practical
purposes they have to be provided for access and maintenance.
The apertures have to be reinforced so that the integrity of the structure is maintained.
Once the aircraft doors are closed and all the hatches and access panels are fitted,
the fuselage is, to all intents and purposes, a monocoque structure.
As aircraft became larger and the air loads greater, the pure monocoque structure was
found not to be strong enough. Additional structural members known as long rungs were
added to run lengthwise along the fuselage, joining the frames together. The light alloy
skin is then attached to the frames and long runs by riveting or adhesive bonding. Long
runs stiffen the skin and assist the sheet materials to carry loads along their length.
This gives us the semi-monocoque type of construction. This is the most common type of construction
used in modern light aircraft.
When cut-outs are made to stressed skin structures, for example to provide access panels, doors
or windows, or when repairs are required to damaged areas, reinforcement, in the form
of doublers or backing plates, is required around the cut-out.
Bulkheads are set in place to separate the different sections of the aircraft, for instance
the engine compartment from the passenger compartment. The bulkhead has the same basic
shape as the frame or former, but almost completely isolates one compartment from the
other. Having said that, some holes will have to be made in a bulkhead in order to allow
control fittings, pipework and electrical cables to pass through.
Bulkheads are usually much more substantially built than the frames, because they are subject
to greater loads. The bulkhead which separates the engine from the passenger compartment
serves to prevent the passage of fire from the engine rearwards, and will do so for some
time, although not indefinitely.
The wings support the weight of the aircraft in the air, and so must have sufficient strength
and stiffness to be able to do this. Their strength and stiffness are determined by the
thickness of the wing, with the thickness and type of construction used being dependent
on the speed requirements of the aircraft. The types of construction are biplane, braced
monoplane and cantilever monoplane.
Very few biplanes fly at more than 200 knots in level flight, so the air loads are low,
which means that a truss-type design which is covered in fabric is satisfactory. The
aircraft has a lattice girder of great rigidity, which is highly resistant to bending and twisting.
Unfortunately, it also generates a relatively large amount of drag, which accounts for the
fact that biplanes fly relatively slowly.
This type of design is also used on low-speed aircraft, although it's very popular with
manufacturers such as Cessna, whose aircraft can hardly be said to be slow. One of the
most famous braced monoplanes was the Spirit of St. Louis that Charles Lindbergh piloted
in his epic solo transatlantic flight.
Let's now look at cantilever wings. The main planes have to absorb the stresses due to
lift and drag in flight and, if of cantilever design, their own weight when on the ground.
This is achieved by building the wing around one or more main load-bearing members known
as spars, which are constructed so that they will absorb downwards bending stresses when
on the ground and the upwards, rearwards and twisting stresses when in flight.
The major structural components of the wings are generally manufactured from aluminium
alloys, with composite materials such as grass-reinforced plastic, carbon-reinforced plastic and honeycomb
structures being used for fairings, control surfaces and flaps, etc.
Bending stress relief is also provided by positioning the major fuel tanks within the
wing. Main planes may be of single spar, twin spar or multi-spar construction. A conventional
structure would consist of front and rear spars, the metal skin attached to the spar
to form a torsion box.
The skin takes the loads due to differences in air pressure and the mass and inertia of
the fuel in the wing tanks. It generates direct stresses in a spanwise direction as a response
to bending moments and also reacts against twisting.
Stringers are spanwise members which give the wing rigidity by stiffening the skin in
compression.
Formers, or ribs as they are sometimes called, maintain the aerophore shape of the wings,
support the spars, stringers and skin against buckling and pass concentrated loads from
engines, landing gear and control surfaces into the skin and spars.
The tail unit, or stabilising unit, sometimes called the empennage, comes in many different
designs. It can be conventional, T-tail, H-tail or V-tail. Some aircraft have their longitudinal
stability and control provided by four planes, canards. The horizontal surfaces which are
known as the tail plane or horizontal stabiliser provide longitudinal stability by generating
upwards or downwards forces as required. The vertical surfaces, the vertical stabiliser
or fin, generate sideways forces as required. Longitudinal control is provided by either
the elevators or the moving tail plane, depending on design. Directional control is provided
by the rudder. Both the tail plane and the fin are subject to both bending and torsional
stresses.
Structurally, the tail unit components are generally smaller versions of the main planes
in that they use spars, ribs, stringers and skin in their construction. On some aircraft
they may also be sealed to provide fuel storage space. They also use the same basic materials.
For instance, aluminium alloys or composites with honeycomb structures. High density expanding
foam is sometimes used within control surfaces to provide greater stiffness at lower weight.
For steady flight the aircraft must be in a state of balance, zero moments around the
axes, and the flying controls enable this to be achieved for all possible configurations
and centre of gravity positions. Secondly, the controls will be required to manoeuvre
the aircraft around its three axes.
The principle of operation of all of the flying controls is explained in the subject Principles
of Flight. Here in this lesson we only consider them as part of the aircraft structure.
Rotation around the longitudinal axis is rolling and is controlled by the ailerons.
Rotation around the lateral axis is pitching and is controlled by the elevators or by a
moving tail plane.
Rotation around the normal axis is yawing and is controlled by the rudder.
On some aircraft rotation around two of the axes may be achieved with one control surface.
For instance, the elevon, which is a mixture of elevator and aileron, is used on tailless
aircraft to give both pitching and rolling. The ruddervator, which is the name of the
control surfaces on a V-tail aircraft, enables both pitching and yawing.
The stabilator is a movable tail plane combining the dual function of horizontal stabiliser
and elevator, that is, it gives both longitudinal stability and control.
The movement of the flying control surfaces in response to the movement of the cockpit
controls in light aircraft is achieved mechanically. This means that the control surfaces are connected
directly to the cockpit controls by a system of cables, rods, levers and chains.
Pitch control can be obtained through the use of either elevators, flying tail planes
or canard control. For the purpose of this lesson, we'll assume that our aircraft has
elevators fitted to the tail plane. If that is the case, then rearward movement of the
control column causes upward movement of the elevator, which causes the aircraft to pitch
nose upwards, and vice versa.
Control in roll is achieved by ailerons. Turning the control wheel to the right causes the
right aileron to move up and the left aileron to move down, giving roll to the right.
Control in yaw is enabled by the rudder. Moving the right rudder pedal forward causes the
rudder to move to the right and causes the aircraft to yaw to the right.
The amount of movement of each control surface to either side of its neutral position is
laid down so that the required control can be achieved over the full range of operating
conditions. The movement is not necessarily the same each side of neutral. For example,
an elevator usually has a greater deflection upwards than downwards.
The limit of movement of the control surface is determined by a mechanical stop. The function
of mechanical stops is to prevent excessive control surface deflection, which may cause
the aircraft structure to be overstressed during normal operations.
A stop which limits the movement to the control surface is called a primary stop. A stop which
limits the movement of the control column or rudder pedals is called a secondary stop.
When the primary stop is closed there will be a small clearance at the secondary stop.
Flutter is the rapid and uncontrolled oscillation of a flight control or the surface to which
it is attached, which occurs as a result of an unbalanced surface. Flutter is caused by
the interaction of aerodynamic forces, inertia forces and the elastic properties of the surface
or structure, and can lead to the catastrophic failure of the structure. Flutter can be prevented
by mass balancing control surfaces to move the control surface centre of gravity closer
to the hinge, and therefore alter the period of vibration.
Flutter must not occur within the normal flight operating envelope of the aircraft. The fact
that an aircraft is poorly maintained, particularly in respect to excessive control surface backlash
or flexibility, may mean that flutter will occur at speeds below the limit airspeed.
An aircraft is trimmed when it will maintain its attitude and speed without the pilot having
to apply any force to the cockpit controls. If it is necessary for a control surface to
be deflected to maintain balance of the aircraft, the pilot will need to apply a force of the
cockpit control to hold the surface in its deflected position. This force may be reduced
to zero by operation of the trim controls. The aircraft may need to be trimmed in pitch
as a result of changes of speed, changes of power or varying centre of gravity positions.
Trimming in yaw will be needed as a result of changes in propeller torque, or if there
is asymmetric power on a twin-engined aircraft. Trimming in roll is less likely to be needed,
but would be required if there was a lateral displacement of the centre of gravity, for
example if the contents of the fuel tanks in each wing were allowed to become wildly
different or, in the case of a twin-engined aircraft, if one engine had failed.
Aircraft are fitted with high-lift devices to reduce the take-off and landing distances.
This permits operation at greater weights from given runway lengths and enables greater
payloads to be carried.
A flap is a hinged portion of the trailing edge which can be deflected downwards and
so produce an increase of camber. There are various types of flap design which all increase
both lift and drag in varying amounts. The most popular type of flap used on light aircraft
is the plane or camber flap. This type of flap increases the camber of the wing when
extended and in doing so can increase the maximum lift of the wing by about 50%.
Lowering flap will increase both the lift and the drag of the aircraft, but not in the
same proportion. Selecting flap will cause a pitch change. The overall change of pitch
will be influenced by the type of flap, the position of the wing and the position of the
tail plane.
Attention or tensile load is one which tends to stretch a structural member. Components
designed to resist tensile loads are known as ties.
Compressive loads are the opposite of tensile loads and tend to shorten structural members.
Components designed to resist compressive loads are known as struts.
Shear is a force which tends to slide one face of the material over an adjacent face.
Riveted joints are designed to resist shear forces.
Bending of the structure involves the three basic loadings. They are tension as the outer
edge stretches, compression as the inner edge squeezes together and shear across the structure
as the forces try to split it. Torsion or twisting forces produce tension at the outer
edge, compression at the inner edge and shear across the structure.
Stress is the internal force inside a structural member which resists an externally applied
force and therefore a tensile load or force will set up a tensile stress, compression
loads and compressive stresses etc.
Stress is defined as the force per unit area
and is measured in units of Newtons per square millimetre.
When an external force of sufficient magnitude acts on a structure the structural dimensions
change. This change is known as strain and is the ratio of the change in length to the
loaded structure.
Buckling occurs to thin sheet materials when they are subjected to end loads and to ties
if subjected to compressive forces.
The maximum load that the designer would expect an airframe or component to experience in
service is termed the design limit load.
The design ultimate load is the design limit load times the safety factor. The minimum
safety factor specified in design requirements is 1.5. The structure must withstand design
ultimate load without collapse. The safety factor is the ratio of the design ultimate
load to the design limit load.
When an aircraft is parked in the open strong or gusty winds could blow the controls about
against their stops with sufficient force to cause mechanical damage. To prevent this
occurring control locks are fitted. These may be external as shown here in the left
and centre pictures. If the control locks are controlled from the cockpit they may be
arranged so that it is impossible to open the throttle until they are removed. Remember
control locks must be removed before flight.
Hand held portable water or water glycol fire extinguishers are used in aircraft. They have
red containers and can be used in passenger cabins for combating fires involving domestic
materials such as paper and furnishing materials.
Water or water glycol fire extinguishers must not be used on fires which involve electrical
equipment or liquids. The glycol is an antifreeze agent which permits operation of the
extinguishers at temperatures as low as minus 20 degrees Celsius. The same type of
extinguisher is also used for ground based applications.
Dry powder extinguishers have a blue or red container with a blue label. The use of dry
powder in crew compartments or passenger cabins of pressurised aircraft is not permitted
although some light aircraft may have this type fitted. Their use inside the aircraft
should be avoided if at all possible as visibility will be restricted by the powder which
can also render inoperative otherwise serviceable electrical equipment. The agent,
potassium bicarbonate, is a non-toxic powder similar to talcum powder. It is very
effective against fires involving flammable liquids, wood, fabric and paper.
It is best known as being the safest extinguisher to use against wheel and brake fires
because it is least likely to cause overheated brake discs to explode but it should not be
used on electrical fires for the reasons previously stated. Dry powder extinguishers are
mostly used for ground based applications.
This extinguishment is stored in signal red, purple, brown or green containers. It is very
effective against electrical and flammable liquid fires. It is only slightly toxic, it
is odorless, non-corrosive and evaporates rapidly needing no residue. It doesn't freeze
or cause cold burns and won't harm fabrics, metals or other materials it contacts. It
acts rapidly on fires by producing a heavy blanketing mist which eliminates air from
the fire source but more importantly it interferes chemically with the combustion process.
It has outstanding properties in preventing reflash after the fire has been extinguished.
Handheld containers using BCF are used in many aircraft and, although it is slightly
toxic, it is quite safe to use provided the cockpit is ventilated subsequent to the fire
being extinguished.
Carbon dioxide is used in trolley based extinguishers on the airfield apron. It is stored in black
or red containers with a black label. It is non-corrosive and extinguishes the flame by
dissipating the oxygen in the immediate vicinity. From a standpoint of toxicity and corrosion
it is the safest agent to use and for many years was the most widely used. Having said
that, if handled improperly it can cause mental confusion and suffocation. It requires
a stronger container than most other agents due to its variation in vapour pressure with
changes of temperature. The use of this agent in aircraft is not permitted. Carbon dioxide
may be used against most fires and is particularly useful against engine fires as it will extinguish
Foam is the principal extinguishment for use on flammable liquid fires. It blankets the
flames by excluding oxygen.
There are numerous types of life jacket or life preserver and the information given here
is of a general nature and doesn't apply to any particular make, model or type. Life jackets
are designed as lightweight items of equipment and should be treated with care at all times.
They are normally stowed in special packs or containers for ease of handling and protection.
Stowage in this manner will ensure that the jacket is maintained correctly folded to ensure
easy and rapid fitting if required for use. If there is any evidence of mishandling or
immersion in seawater life jackets should be rejected for operational use. Instructions
for fitting are printed on the container and or the jacket.
All the life jackets are basically similar in design. Buoyancy is achieved by inflating
the jacket with carbon dioxide gas which is stored under pressure in a small cylinder
and released manually by the operation of a red toggle or lever. Once it has been released
the gas will pass through a non-return valve into the jacket and cannot be stopped. The
life jacket can be inflated or topped up orally. The manual inflation tube usually contains
a valve which can be operated to lower or release the pressure in the jacket. The life
jacket is coloured either brilliant yellow or flame red as an aid to location and may
contain various equipment to expedite the location of the individual wearing it.
If you are flying in a light aircraft over extensive areas of water the recommended practice
in respect of life jackets is that they should be worn uninflated for the flight. If an emergency
occurs and the aircraft appears to be going to ditch the pilot will not appreciate his
passengers struggling to don their life jackets in close proximity to him in the cramped confines
of a light aircraft cabin while he is trying to concentrate on much more serious matters.
In the unfortunate event that the aircraft does ditch life jackets should not be inflated
until the wearer has left the aircraft.
Although the ground crew will have checked the aircraft thoroughly it's up to the pilot
to ensure that it is safe for flight. To this end a series of checks is carried out prior
to, during and after the flight.
Initially students will have to rely on the expanded checklist to carry out these checks.
As students gain experience the checks should become instinctive and the time taken to do
them will decrease but every so often students should compare their actions against the expanded
checklist.
Carrying out the external pre-flight check will inevitably bring the student close to
No part of the student's body or their equipment should be allowed to come within the arc that
would be covered by the revolving propeller.
After time students should be able to locate items in the cockpit without looking.
The student's spare time should be used to practice checks but they should beware of
learning them parrot fashion. Every check should be made to mean something and when
appropriate the checklist should always be used.
The danger of loose articles in aircraft cockpits cannot be overstressed. Numerous accidents
have been attributed to foreign objects left in the aircraft by untidy pilots. Personal
items such as pens, pencils, keys, combs and coins can all be the cause of accidents if
they become tangled in the flying controls.
If you do drop anything in the cockpit then you must be strenuous in your efforts to retrieve
it as soon as possible. If you are unable to locate it then it must be reported to the
ground crew who will make a thorough search until it is found. Only after it is found
will the aircraft be allowed to fly again.
On the apron there are other aircraft and items of ground equipment and personnel, for
example refuelling crews. As a consequence of this all pilots must be conscious of the
safety precautions required when starting engines and when taxiing onto or away from
the apron. Extreme caution must be exercised and the taxiing speed must be kept to a minimum.
Remember the pilot is always responsible for the safety of the aircraft.
Just as the checks must become well practised so must the initial reactions to emergencies.
Emergencies are fortunately rare in modern aircraft, but pilots must be able to deal
promptly with any situation that may arise. The emergency drills should be practised frequently
in the air and the practices should be as realistic as possible. Before being sent solo
the student must know what actions to carry out in the event of a fire, both on the ground
and in the air, an engine failure after take-off and a forced landing.
The functions of the landing gear are firstly to provide a means of manoeuvring the aircraft
on the ground, secondly to support the aircraft at a convenient height to give clearance for
propellers and flaps, etc., and thirdly to absorb the kinetic energy of landing.
Once airborne the landing gear serves no useful purpose and is dead weight. It would be ideal
to replace it with some ground-based equipment which detached on take-off, as with this Messerschmitt
Comet. But while it's possible to achieve manoeuvring and support for the aircraft with
ground-based equipment, no satisfactory alternative exists for absorbing the kinetic energy of
landing and providing a means of controlling deceleration. For this reason a vast amount
of research has gone into the design of undercarriage units in order to reduce their weight and
their stowed volume when retracted.
With slow, light aircraft and some larger aircraft on which simplicity is of prime importance,
a non-retractable landing gear is often fitted. The reduced performance caused by the drag
of the landing gear during flight is offset by its simplicity, its reduced maintenance
and also its low initial cost. With higher performance aircraft drag becomes progressively
more important and the landing gear is retracted into the wings or fuselage during flight.
There are, however, penalties of increased weight, greater complication and additional maintenance.
There are three main types of fixed landing gear. Those which have a spring steel leg,
those which employ rubber cord to absorb shocks and those which have an oleonimatic strut
to absorb shocks. Exceptions include aircraft with rubber in compression, spring coil and
liquid spring struts.
Spring steel legs are usually employed at the main undercarriage positions. The leg
consists of a tube or strip of tapered spring steel, the upper end being attached by bolts
to the fuselage and the lower end terminating in an axle on which the wheel and brake are assembled.
When rubber cord is used as a shock absorber the undercarriage is usually in the form of
tubular struts, designed and installed so that the landing force is directed against
a number of turns of rubber in the form of a grommet or loop.
Some fixed main undercarriages and most fixed nose undercarriages are fitted with an oleonimatic
shock absorber strut. The design of individual struts varies considerably but one point worthy
of consideration is the fitting of spats to oleonimatic struts.
Spats are an aerodynamic fairing which may be required to minimise the drag of the landing
gear structure. One drawback to their use is the problem of them picking up mud when
landing or taking off from grass airfields. This can add considerably to the weight of
the aircraft and may affect take-off performance. To avoid this eventuality, if any mud has
been picked up, the spats must be removed, cleaned and replaced before the next take-off.
The nose undercarriage is usually a lighter structure than the main gear units since it
carries less weight and is usually subject only to direct compression loads. The nose
wheel must be able to castor freely. Castoring is the ability of the nose wheel to turn to
either side in response to differential braking.
A method of steering is required to enable the pilot to manoeuvre the aircraft safely
on the ground. Early methods involved the use of differential braking but today the
nose wheel steering of most light aircraft is controlled by the operation of the rudder
pedals which are connected to the nose wheel by control rods or cables.
Due to the flexibility of tyre sidewalls, an unstable rapid sinusoidal oscillation or
vibration known as shimmy can be induced into the steerable parts of the undercarriage.
Excessive shimmy, especially at high speeds, can set up vibrations throughout the aircraft
and can be dangerous. Worn wheel bearings and uneven tyre pressures can both increase
the tendency of the wheels to shimmy. Shimmy can be reduced in several ways, for instance
by fitting a shimmy damper or by having heavy self-centring springs fitted in the nose wheel
steering control rod run. Some larger aircraft have double nose wheels fitted while twin
contact wheels have also been found to be effective in minimising tail wheel shimmy.
An increasing number of light aircraft are fitted with a retractable landing gear for
the purpose of improving aircraft performance. Retraction is normally affected by a hydraulic
system but pneumatic or electrical systems are also used. In some instances power is
used for retraction only, extension being affected by gravity and the slipstream. Retractable
landing gear is also provided with mechanical locks to ensure that each undercarriage is
locked securely in the retracted and extended positions, devices to indicate to the crew
the position of each undercarriage and a means by which the landing gear can be extended
in the event of a failure of the power source. In addition a system is provided to prevent
retraction of the landing gear with the aircraft on the ground and to guard against landing
with the landing gear retracted. Undercarriage wells are normally sealed by doors for aerodynamic
reasons.
Aircraft wheels are fitted with pneumatic tyres. These can consist of both a rubber
inner tube and an outer cover. The tube contains compressed air which absorbs shock and supports
the weight of the aircraft, while the cover restrains and protects the tube from damage,
maintains the shape of the tyre, transmits braking effort and provides a wearing surface.
The tyre cover consists of a casing made of rubber which is reinforced with plies of cotton,
rayon or nylon cords. The cords are not woven but arranged parallel in single layers and
held together by a thin film of rubber which prevents cords of adjacent plies from cutting
one another as the tyre flexes in use. During construction of the cover the plies are fitted
in pairs. Each pair is termed a pocket which is set so that the cords of adjacent plies
are at 90 degrees to one another. This is called cross-ply. The tyre manufacturers give
each tyre a ply rating. This rating doesn't relate directly to the number of plies in
the tyre but is instead the index of the strength of the tyre. These special plies
are termed breaker strips to absorb and distribute load shocks and also to protect the casing
from concussion damage. Narrow plies embedded in thick layers of rubber are situated between
the casing and the tread. The casing is retained on the rim of the wheel by interlocking the
plies around inextensible steel wire coils to form ply overlaps. This portion of the
cover is known as the bead.
To assist you, should you ever need to describe the cover, it's divided into regions or sections
as illustrated here. The tread of the tyre is situated on the crown and shoulder section.
The sidewall allows the tyre to be flexible and the bead is seated inside the flange of
the wheel. Note that the term tread is applied irrespective of whether the rubber is plain
and smooth or moulded on a block pattern. The most popular tread pattern is that termed
ribbed. It's formed from circumferential grooves around the tyre.
The size of a tyre is marked on its sidewall and includes the outside diameter, the inside
diameter and width of the tyre, all in inches. The ply rating, the index of the tyre's strength,
is also marked on the sidewall. Normally it's shown as an abbreviation, for instance 16
PR, but occasionally it's shown in full as 16 ply rating.
The speed rating of the tyre, the maximum speed for which it's designed, is imprinted
on a panel on the sidewall of some high speed tyres. The rating takes account of pressure
altitude, ambient temperature and wind component, enabling the maximum take-off weight the tyres
can sustain to be calculated. Green or grey dots painted on the sidewall of the tyre indicate
the position of the all vents. All vents prevent pressure being trapped between the
plies, which would cause disruption of the tyre carcass if it was exposed to the low
pressures experienced during high altitude flight.
A red dot or triangle indicates the lightest part of the tyre. If this is placed opposite
the valve during tyre fitting then it assists in balancing the wheel assembly.
Tyres must be protected from excessive heat, dampness, bright sunlight, contact with oil,
fuel and hydraulic fluid. All of these have a harmful effect on rubber.
Covers should be placed over the tyres when the aircraft is to be parked for any length
of time or during the periods when oil, fuel, cooling or hydraulic systems are being drained
or replenished. Any fluid inadvertently spilt or allowed to drip onto a tyre must be wiped
off immediately.
When tyres are first fitted to a wheel they tend to move slightly around the rim. This
phenomenon is called creep and at this stage it is considered normal.
After the tyre has settled down however this movement should cease.
In service the tyre may tend to continue to creep around the wheel. If this creep is excessive
on a tyre fitted with an inner tube it will tear out the inflation valve and cause the
tyre to burst.
Creep is less of a problem with tubeless tyres as long as the tyre bead is undamaged and
any pressure drop is within limits. Creep is less likely to occur if the tyre air pressure
is correctly maintained.
To assist in this tyre manufacturers specify a rated inflation pressure for each tyre.
Witness marks called creep marks are painted both on the bead of the tyre and on the flange
of the wheel.
When the tyre is first fitted to the wheel the creep marks will be fully aligned with
each other as shown here and in a perfect world would remain that way for the life of
the tyre. However for a variety of reasons including heavier than you would like landings,
harsh braking and low tyre pressures the tyre will move around the wheel and the creep marks
will move in relation to each other. Movement of the tyre around the wheel must not be such
that the creep marks become fully misaligned with each other as illustrated here.
During your pre-flight inspection tyre covers must be examined for cuts, bulges, embedded
stones, metal or glass, signs of wear, creep etc. Embedded stones, metal, glass etc. must
be removed and the cuts probed by a technician to ascertain their depth. Cuts in the tyre
cover penetrating to the cords render the tyre unserviceable. The cords will be recognisable
as pieces of white fibrous material embedded in the rubber of the tyre. Whether or not
the tyre can either be repaired or must be renewed is governed by the depth of the cut.
Bulges may indicate a partial failure of the casing. If the fabric is fractured then the
tyre must be renewed. Any flat spots on the tyre tread, especially those caused by skidding
or aquaplaning, must be examined with great care to determine if the fabric of the tyre
has been weakened unduly. Damage caused by aquaplaning will make the rubber of the tyre
look as if it has been melted and in some cases gives it the appearance of having been
bleached. If there is any doubt of the serviceability or otherwise of a tyre then it should be rejected.
Aquaplaning is a phenomenon caused by a wedge of water building up under the tread of the
tyre and breaking its contact with the ground. Aquaplaning speed, measured in nautical miles
per hour, is the speed that the tyre loses contact with the ground. It can be found by
applying the formula aquaplaning speed equals 9 times the square root of P, where P is the
tyre pressure in pounds per square inch. Or aquaplaning speed equals 34 times the square
root of P, where P equals the tyre pressure in kilograms per centimetre squared.
The possibility of aquaplaning increases as the depth of the tread is reduced. It's therefore
important that the amount of tread remaining is accurately assessed.
In common with most braking systems, aircraft wheel brakes function by using friction between
a fixed surface and a moving one to bring an aircraft to rest, converting kinetic energy
into heat energy. The amount of heat generated in stopping even a light aircraft is extremely
large, and the bigger the aircraft the greater the heat. Although it doesn't exactly come
within the scope of this lesson, it's worth noting that the amount of heat absorbed by
the brakes of a jumbo jet during an aborted take-off is sufficient to put an elephant
into orbit and cook it. The problem of dissipating this heat has been a challenge to aircraft
designers and scientists for years.
Most light aircraft now use hydraulically activated disc brakes as their means of slowing
down or stopping. These use a series of fixed friction pads bearing on or gripping a rotating
disc, similar in principle to the disc brakes on a car. The friction pads are made of an
inorganic friction material and the discs are of forged steel with a specially case-hardened
surface. It's this surface and interior structure combination which causes the plates to explode
if doused with a liquid fire extinguishment when they are red hot. In the unfortunate
event of a wheel or brake fire, the best extinguishment to use is dry powder.
If pressure is applied to the foot brake pedals or the handbrake lever, hydraulic pressure
will build up in the slave cylinder behind the piston. This pressure will cause the piston
to move over within the caliper unit, pushing the brake pad against the brake disc.
The reaction of the right hand brake pad pushing against the disc will cause the caliper
unit to move in the opposite direction to the piston, carrying the left hand brake pad
with it until the disc is squeezed between the two pads.
The force applied to the brake pads will be proportional to the effort applied to the
tow brakes. The tow brakes can be applied together or, if differential braking is required
for manoeuvring the aircraft within tight spaces, the tow brakes can be used separately.
Most light aircraft have a hand-operated parking brake. When selected, the handbrake
will lock the hydraulic pressure applied by the foot brakes in the brake system.
Before flight the brake system should be checked for any signs of malfunction. It will be relatively
easy to check for signs of hydraulic fluid leaks. Any fluid leaking from the slave cylinders
within the caliper units will either drip straight onto the ground or dribble down the
sidewall of the tyre. In either case it should be fairly easy to see.
Making sure before you touch it that the brake assembly is not still hot from previous use,
check that the caliper unit is firmly attached to the gear, and while you're in the vicinity
have a look at the wear on the brake pads.
If you can, make sure the part of the disc upon which the brakes come into contact is
bright and shiny. Any rust or pitting on the braking surface of the disc will reduce
its effectiveness and may be indicative of a malfunction of the system. Any problems
associated with the braking system should be rectified before the aircraft is moved.
The accepted definition of the term piston engine is an engine in which the working fluid
is expanded in a cylinder against a reciprocating piston. This definition could be applied to
any reciprocating type of internal combustion engine, but here we'll use it to imply the
type of engine used in most light aircraft, namely the four-stroke gasoline spark ignition engine.
The piston engine converts the heat energy of the fuel it burns into mechanical energy.
This conversion is accomplished by introducing an air-fuel mixture into a cylinder which
is then compressed by a piston. The mixture is then ignited by an electric spark.
A subsequent combustion causes a rapid rise in the temperature and pressure of the gases
within the cylinder, which forces the piston downwards. This linear piston movement is
converted into rotary motion by the crankshaft.
When all of the useful pressure has been used, the burnt gases are exhausted to atmosphere.
The principle of operation of all four-stroke internal combustion engines is based upon
the Otto cycle, originally propounded by Bode Roscha in 1876. However, Otto was the
first to succeed in producing a working engine using this cycle.
The four-stroke cycle involves four strokes of the piston and generates two revolutions
of the crankshaft.
The stroke of the piston is the extent of its travel within the cylinder. The lower
limit is called bottom dead centre or BDC and the upper limit is called top dead centre
or TDC.
The four strokes are called induction, compression, power and exhaust.
We'll use a simple single-cylinder engine to illustrate what happens in each of the
four strokes.
The timing diagram for the Otto cycle is shown on the right side of the screen.
The induction stroke. At top dead centre the inlet valve opens and the piston descends
creating a depression, drawing the air-fuel mixture into the cylinder.
At bottom dead centre the inlet valve closes.
The compression stroke. With both valves closed the piston ascends and the combustible mixture
is compressed into the combustion chamber, which is the space formed in the cylinder
above the piston when it's at top dead centre.
The power stroke. A spark plug ignites the mixture and combustion causes a rapid rise
in the temperature and pressure. Because both valves remain closed the increased gas pressure
forces the piston down.
The exhaust stroke. The exhaust valve now opens as the piston ascends and the burnt
gases, which are now at a very reduced pressure, escape to atmosphere. The piston continues
to rise to top dead centre, where the cycle begins again.
In practice the Otto cycle proved to be inefficient and the times of valve openings and closings
and the ignition point had to be modified. A practical timing diagram will help to illustrate
the reasons for the modifications.
Opening the inlet valve before top dead centre ensures that it is fully open early in the
induction stroke. The time lag, which would normally occur between the piston moving down
and the mixture flowing into the cylinder because of the inertia of the mixture, is
eliminated. The incoming mixture can thus keep up with the descending piston.
The momentum of the incoming mixture increases as the induction stroke proceeds and towards
the end of the stroke is such that the gases will continue to flow into the cylinder even
though the piston has passed bottom dead centre and is moving slightly upwards. Closure of
the inlet valve is therefore delayed until after bottom dead centre, when the gas pressure
in the cylinder approximately equals the gas pressure in the inlet manifold.
As the piston continues to move upwards the inlet valve closes and the gases are compressed.
By squeezing the gas into a smaller space the pressure it will exert when burnt is proportionally
increased.
Before the piston reaches top dead centre on the compression stroke the mixture is ignited
by a spark. The piston is carried past top dead centre by the momentum of the moving
parts while the flame is spreading. The intense heat raises the pressure rapidly to a peak
value which is only reached when combustion is complete. By timing the ignition point
complete combustion, and therefore peak pressure, is arranged to coincide with the crankshaft
being 10 degrees after top dead centre. This allows the engine to gain maximum advantage
from the combustion.
Most of the energy in the gases has been converted into mechanical energy by the time the piston
has moved halfway down the cylinder. If the exhaust valve is opened around this point
any residual gas pressure will start the first phase of exhaust scavenging. Opening the exhaust
valve before bottom dead centre will also ensure that the pressure of gases remaining
in the cylinder does not resist the upward movement of the piston in the exhaust stroke.
Finally the piston moves upwards forcing the remaining gases out of the cylinder. The exhaust
valve is left open after top dead centre to permit the gases to scavenge the cylinder
as completely as possible by their momentum.
The core around which the engine is built is called the engine block. It's usually made
from steel alloy. Depending on the manufacturer of the engine cylinders are bored into or
fitted onto the engine block. The cylinders are made of alloy steel. They resist the pressure
of combustion and provide a working surface for the piston rings. If the engine is air
cooled then the cylinders are finned to increase the cooling area. If the engine is liquid
cooled a large one piece jacket houses cylinder liners to provide a cooling liquid container
around them.
Pistons made from light alloy are fitted within the cylinders. Around each piston are
fitted two or in some cases three piston rings. The piston rings are made from cast iron and
they ensure that no gases leak from above the piston into the crankcase where the oil
is usually kept.
The connecting rod transmits the forces of combustion to the crankshaft which converts
the linear motion of the piston into rotary motion at the crankshaft. The connecting rod
is made from H section high tensile steel which combines lightness with the strength
necessary to withstand the compressive and tensile loads imposed as the piston changes
direction. The connecting rod is joined to the piston by a gudgeon pin which fits through
the small end of the rod and is joined to the crankshaft by a large circular bearing
at its big end.
The cylinder head is generally made of aluminium alloy and is finned to improve heat dissipation.
It seals one end of the cylinder to provide a combustion chamber for the mixture. The
cylinder head accommodates the valves and the sparking plugs and supports the valve
rocker arms.
The valve operating gear consists of a camshaft or shafts driven by the crankshaft at half
crankshaft speed. A camshaft is driven at half crankshaft speed because each valve is
only required to open and close once per working cycle, that is once every two revolutions
of the crankshaft.
An accessory housing is a casing mounted at the rear of the block. It encloses the drive
gear trains for the camshafts, the fuel, oil, pneumatic and vacuum pumps, electric generator,
magnetos and tachometer.
During the four strokes the changing pressures within the cylinder can be measured and indicated
graphically on a device which produces an indicator diagram. This device is only used
during the development stage of an engine. The illustration shows an indicator diagram
opened out so that the pressure areas under the curve can more easily be compared and
measured.
The area within the power column represents work done on the piston during the power stroke.
The blue areas represent work done by the piston in compressing the gases and exhausting
the cylinder against exhaust back pressure, that is the pressure which the atmosphere
exerts against the outflowing gases.
Subtracting the sum of the blue coloured areas from the area within the power column leaves
the red block.
The red block is indicative of the average pressure during the working cycle. This is
termed the indicated mean effective pressure or IMEP.
We can use the indicated mean effective pressure in a formula which will allow us to determine
the indicated horsepower of our particular engine.
In the formula the P represents the indicated mean effective pressure in pounds per square
inch. However, we are still required to determine several other values.
The L represents the length of the stroke in feet, the length of the stroke being the
distance the piston moves between top dead centre and bottom dead centre.
The A will be the area of the cylinder in square inches.
N is the number of cylinders that the engine has.
E is the effective working strokes per minute or the RPM of the engine.
And 33,000 is a constant used to change foot pounds of work per minute into horsepower.
One horsepower equals 33,000 foot pounds per minute or 550 foot pounds per second.
Bearing in mind that only one function of the formula is variable, that is the RPM of
the engine, then it is logical to assume that engine power output is directly proportional
to RPM. Increase the engine speed and the power output increases. Decrease the engine
speed and the power output falls.
The compression ratio of an engine is the ratio of its total volume enclosed in the
cylinder with a piston at bottom dead centre
to the volume remaining at the end of the compression stroke with the piston at top
dead centre.
The total volume of the engine is the sum of the swept volume which is the volume displaced
by the piston as it moves from bottom dead centre to top dead centre and the clearance
volume.
An example will help to clarify the term. If the swept volume of an engine is 1800 cubic
centimetres and its clearance volume is 300 cubic centimetres, what is the compression
ratio of the engine?
First we must add the swept and clearance volumes to obtain the total volume, so 1800
plus 300 is 2100.
Then we must divide the total volume by the clearance volume, that is 2100 divided by
300 and the answer is 7.
So the compression ratio of this engine is 7 to 1.
The piston engine is a heat engine. Its purpose is to convert the energy released from burning
fuel into mechanical energy and by so doing generate power that we can use.
The thermal efficiency of a light aircraft piston gasoline engine may reach 30%.
This means that 70% of the heat energy released by the fuel is wasted.
Around 40% will leave with the exhaust gas but 30% stays behind to raise the temperature
of the engine components and the oil.
If nothing was done to alleviate this situation then several problems could occur.
Engine components could start to fail structurally. The oil could break down and lose its lubricating
properties.
The fuel could ignite as soon as it entered the cylinder before the spark plug fires.
This is called pre-ignition.
Pre-ignition causes loss of power and may exacerbate the overheating problem.
The incoming mixture may get too hot which will cause unstable combustion, usually called
detonation or pinking.
This too will cause loss of power and further overheating.
So far we've only considered what would happen if the engine overheated.
Let's now look at the problems of over-cooling.
Maximum thermal efficiency is only achieved at high engine temperatures.
Low engine temperature increases the viscosity, the internal friction of the oil which will
reduce the brake horsepower.
To burn the fuel must vaporize. Very low engine temperatures would prevent complete vaporization
and some of the fuel would not be burnt, reducing the power output of the engine.
To work efficiently the engine needs to be at the highest temperature consistent with safe operation.
To maintain the engine temperature within a range which will satisfy its working requirements,
a cooling system is required.
Most light aircraft engines are air-cooled.
The simplicity of the system makes it virtually maintenance-free and it's much lighter than
the alternative, the liquid cooling system.
Having said that, the liquid cooling system is more efficient.
It gives more precise control of engine temperature and generates less drag.
Thus it's used on high-performance engines.
One of the major factors governing the efficiency of an air-cooled system is the air temperature.
Although the ambient air temperature can vary widely with changes in climate and altitude,
it must be borne in mind that dissipation of the heat will be more rapid as air temperature decreases.
The speed of the airflow is the other major factor affecting the efficiency of the air-cooling system.
Uncontrolled, the speed of the airflow over the cylinders will vary with the speed of the aircraft
and therefore cause the temperature of the cylinders to fluctuate accordingly.
Several methods of improving the efficiency of cooling systems can be employed.
Firstly, the engine components are made of materials with a high conductivity.
For instance, cylinder heads are sometimes made from aluminium alloys.
The walls of the cylinders are thinned to increase the cooling area.
The fins are thin in section and may be extended to increase the area at local hotspots
such as the exhaust ports in an attempt to achieve an even temperature throughout.
It must not be forgotten that the oil used to lubricate the engine is also a cooling medium.
The engine is streamlined by surrounding it by a cowling.
This puts it within a controlled environment and also reduces the drag which it would otherwise generate.
Baffles are directional air guides which direct the airflow fully around the cylinder.
They must be close fitting and are sealed against the cowling to ensure that all of the airflow is over
the cylinders
and that each part of the cylinder is supplied with equal amounts of cooling air.
Cow flaps and gills can be fitted to the cowlings to ensure control of the temperature.
They are usually manually controlled on light aircraft
and when open the airflow over them causes a pressure drop which accelerates the airflow within the cowlings.
Prior to engine start the air inlet and outlet to the cowlings must be checked to be clear of obstruction.
The condition of baffles and cowling must be checked.
The operation of the cow flap, if fitted, must be checked.
Remembering that as well as lubricating the engine oil plays an important part in engine cooling, check the oil contents.
In high power settings with low airspeed conditions, such as occur at take-off,
the cow flaps should be selected open to increase the flow rate over the cylinders and so increase cooling.
In the climb and cruise the cow flaps or gills can be adjusted to maintain the engine temperature at the
optimum.
In descent engine power is reduced and less heat is generated.
If the aircraft is placed into a dive this will tend to over-cool the engine and this will lead
to what is known as thermal shock.
Thus the cow flaps should be closed during descent to reduce the cooling effect.
Anticipating a go-around, the cow flaps should be opened on final approach.
During ground running, engine oil temperatures and cylinder head temperatures should be closely monitored to ensure the limitations are not
exceeded.
Although initially after engine start the cow flap should be closed to encourage quick engine warm-up, they should subsequently
be opened to prevent overheating.
During flight, if the cylinder head temperature becomes excessive, there are several methods of cooling the engine.
Obviously, as has just been stated, the cow flaps can be opened.
Perhaps less obviously, the mixture can be richened. The excess fuel will evaporate and cool the charge.
It's also logical to expect that a reduction in engine power will tend to lower engine temperatures and likewise an
increase in forward speed will have the same effect.
Engine components which move against each other are subject to friction and friction causes a loss of engine power and
wear of the components.
Friction can be reduced by inserting a lubricant between the moving surfaces.
In the case of aviation engines, the lubricant is oil.
Pressure lubrication requires that the oil is forced between the moving parts. Alternatively, the components can be splash lubricated.
The primary task of the oil is to reduce friction and wear, but it has a number of secondary functions
which are extremely important.
Perhaps the most important of these secondary tasks is that of cooling. The flow of oil through the engine picks
up heat which is then dissipated in an oil cooler.
As the oil flows through the engine, it carries away the by-products of combustion and thus cleans the engine.
A filter will clean the oil and allow it to return back to the system.
Most of the components of the engine are ferrous in nature and will, if left alone, corrode. The oil prevents
this corrosion.
The oil can act as a hydraulic medium, for instance, when used as the hydraulic fluid in a variable pitch
propeller mechanism.
The oil system can also act as an indicating medium, giving information to the pilot of engine condition and power
output.
There are two lubrication systems in common use. They are the wet sump system and the dry sump system.
Most light non-aerobatic aircraft use the wet sump system, where the oil is stored in the bottom of the
engine, usually called the sump.
This simplifies construction, but leads to a number of disadvantages.
In the wet sump engine, the oil is circulated by the pressure pump. The oil then passes through a high
pressure filter into internal drillings in the crankcase.
These are called the oil gallery. Ducts in the oil gallery transfer the pressure oil to provide lubrication for the
various drives in the accessory casing, and also the main crankshaft bearings.
Annular grooves are cut in the crankshaft bearings, and these are aligned with holes drilled in the crankshaft.
These fill with pressure oil, which is forced through the transverse drillings in the crankshaft into the big end bearings.
Oil droplets, being squeezed out of the bearings, will collide with the crankshaft as it rotates, with the result that
the inside of the crankcase is filled with a fine oil mist.
This will provide splash lubrication for the cylinder walls, the pistons, the small end bearings, gear teeth, and valve mechanisms,
etc.
Lubrication for the valve gear rocker arms is fed through external oil pipelines.
The first disadvantage of the wet sump lubrication system is that there are lubrication difficulties during manoeuvres.
Oil is thrown around in the engine sump, causing parts of the engine to be under-lubricated, and likewise parts
of the engine to be over-lubricated, which can be equally hazardous.
Inverted flight can be particularly dangerous to the health of the engine.
The second disadvantage of the wet sump lubrication system is that the temperature of the oil is more difficult to
control, because it is stored within the engine in the sump, which itself is hot.
A third disadvantage of the wet sump oil system is that the oil becomes oxidised and blackened by the continual
stirring action of the crankshaft and the connecting rods.
Finally, the amount of oil available is limited by the sump capacity.
The dry sump system overcomes all of these problems by storing the oil in a remotely mounted tank.
The tank is made of sheet metal, suitably baffled internally to stop the oil surging around while the aircraft is
manoeuvring.
Whenever possible the tank is placed above the level of the engine to ensure a gravity feed into the engine.
The tank forms a reservoir of oil large enough for the engine s requirements, plus an air space.
The air space is very important. It allows for several situations.
For instance, the expansion of the oil when it gets hot, as it inevitably will.
Also frothing, due to aeration of the oil as it flows around the engine, will occur, requiring a greater volume
of space.
Another situation that requires the air space is caused every time the engine is shut down.
The walls of the crankcase are saturated with oil that drains into the sump.
The oil will remain there until the engine is restarted, when the scavenge pump, which is fitted at the bottom
of the sump, will return it to the oil tank.
Finally, if the engine has any oil operated devices fitted to it, such as a variable pitch propeller, the displacement
that these mechanisms cause has to go somewhere, and the tank serves as the receptacle for it.
The pressure pump consists of two deep-toothed spur gears rotating in a close-fitting pump casing.
It s fitted on the accessory casing and is driven from the crankshaft by gears.
Oil pressure is sensed at the outlet of the pressure pump.
Unless it s controlled, the oil pressure will vary depending on the speed of the pump, the temperature of the
oil and the resistance offered by the engine components.
The capacity of the pump must be sufficient to supply a minimum oil pressure under its most adverse running conditions,
that is, when the engine is running slowly and the oil inlet temperature is very high.
Because of this, under normal running conditions the pressure available would be much greater, possibly great enough to damage the
engine or its components.
To maintain the oil pressure more or less constant during normal engine running conditions, and also to prevent excessive oil
pressure in the system, a pressure relief valve is fitted between the inlet and the outlet of the pump.
When the pressure reaches a predetermined figure, the valve opens and sufficient oil is returned to the inlet side of
the pump to limit the maximum oil pressure.
A coarse wire mesh suction filter is fitted between the tank and the pressure pump. It s designed to remove
large solid particles from the oil before it enters the pressure pump, and thus it prevents damage to the pump.
The pressure filter is fitted downstream of the pressure pump before the oil enters the engine. It s designed to
remove very small solid particles before the oil passes to the bearing surfaces, and thus it prevents damage to the
engine.
The scavenge pump returns the oil which has passed through the engine into the sump back to the oil tank.
Essentially it s the same design as the pressure pump, but larger.
This is because to maintain a dry sump the scavenge pump needs to be of a larger capacity than the
pressure pump.
As has already been mentioned, the oil is used as a cooling medium in the engine. If the oil gets
too hot then it will fail as a lubricant, and as a consequence the engine will also fail.
To prevent the oil temperature rising too high, an oil cooler is introduced into the system.
The oil cooler consists of a matrix block which forces the oil into a thin film as it passes through
it.
The cooler matrix is subjected to a flow of cold slipstream air which has been directed through the cowlings.
Engines of higher power can have shutters fitted to the oil cooler. These are used to control the flow of
air through the cooler. They can be manual or automatic in operation.
Prior to engine start, the engine oil will normally be cold and viscous.
The cooler matrix is fairly delicately constructed and can easily be damaged by high internal pressures.
To prevent damage to the cooler matrix, an anti-surge valve is placed in parallel with it.
This will allow cold, thick oil to bypass the cooler and go straight back to the tank.
Different engines operating in different ambient conditions will need different grades of oil.
The grade of an oil is mainly dictated by its viscosity.
Viscosity is defined as the measure of the internal friction of a fluid.
A liquid that flows freely has a low viscosity and one that flows slowly has a high viscosity.
The viscosity of an oil will change with changes in its temperature.
A cold oil will have a higher viscosity and a hot oil will have a reduced viscosity.
There are various standards employed to determine the viscosity of oils. They provide a standard by which different oils can
be compared.
The two standards which are generally employed in aviation are the Society of Automobile Engineers, or SAE, and the Sabolt
Universal.
Both systems use numbers to denote the viscosity. The lower the number, the thinner the oil.
There are in use oils that have two viscosity values, for instance, SAE 1550.
Oils with two viscosity values are called multigrade oils. They have the characteristics of low viscosity at low temperatures and
high viscosity at high temperatures.
The type of oil normally used in aircraft piston engines is mineral based. However, some engine manufacturers have trialled and
approved the use of semi-synthetic oils.
If an oil contains no additives then it is called a straight oil.
A straight oil can be recognized by the fact that it only has an identification number. For instance, the bottle
marked OIL 80 contains a straight oil.
Note that the oil has a viscosity rating of SAE 40.
Generally speaking, a straight oil is only used when running in new engines.
However, there are certain requirements of some engines that straight oils cannot answer. To satisfy these requirements, additives must be
mixed with the oil.
These additives can take the form of antioxidants, detergents, and oiliness agents.
Oils with additives are called compound oils. A compound oil of the same viscosity as the straight oil shown previously,
SAE 40, would be kept in a container marked W80.
This particular oil is used where specific cleaning qualities are required.
Oil pressure and temperature readings, when interpreted correctly, are indicative of the mechanical integrity of the engine.
A scan of the aircraft instruments must always include these vital items. The information they give may enable the pilot
to take action which will prevent a bad situation getting much worse.
As part of the pre-flight check, the oil contents must be ascertained to be sufficient for the flight we
are to undertake. The oil filler cap must be secure and a thorough check made to ensure that there are
no oil leaks.
Radial and inverted engines can suffer from a condition known as hydraulicing. This happens if oil accumulates between the piston
and the cylinder head in the cylinders at the bottom of the engine.
Because it's incompressible, if oil is trapped between the piston and the cylinder head on the compression stroke, a hydraulic
lock will be formed.
If the starter motor is being used to start the engine, then something will have to give. The piston might
break, the connecting rod might bend, the cylinder may be torn from the crankcase or the crankshaft may break. None
of these is a pleasant prospect.
To prevent hydraulicing, the engine must be rotated by pulling through on the propeller, after confirming that the magnetos are
off before the propeller is touched.
After engines start, the oil pressure must be indicated within a specified time. If the engine is started from cold,
then the oil pressure may initially seem excessively high. However, as long as the oil pressure drops to normal as
the engine warms up, this can be considered acceptable.
If the aircraft you are operating has a dry sump engine, then its oil contents must be checked immediately after
the engine has stopped, realistically within a few minutes of shutdown. This ensures that the tank contents are recorded accurately
before the oil migrates into the sump under the influence of gravity.
If on the other hand the engine has a wet sump, then a period of 15 to 20 minutes must
elapse before the contents are checked.
There are four main indications associated with the failure of the lubrication system. They are oil pressure too high, oil
pressure too low, fluctuating oil pressure, oil pressure falling to zero.
Too high an oil pressure with normal oil temperature may be a function of the pressure relief valve being set
incorrectly, but this is unlikely unless the component has been worked on just prior to the flight.
If on the other hand the high oil pressure is accompanied by a low oil temperature, the high oil pressure
may be caused by the relatively high viscosity of the oil at low temperatures, making it difficult for the oil
to pass through the engine. The oil pressure relief valve may be unable to control the situation completely until the
oil temperature rises.
As was stated before, as long as the oil pressure drops to normal as the engine warms up, this situation
can be considered acceptable. However, if the pressure remains at dangerously high values, the engine must be stopped to prevent
damage.
Too low an oil pressure with normal or even low oil temperature could once again mean that the oil pressure
relief valve has been set incorrectly. But as before, this is unlikely unless the component has been worked on just
prior to the flight.
The problem here is more likely to be excessive clearance on the bearings due to wear or, potentially even worse,
leaks from the output side of the pressure pump. If the minimum oil pressure cannot be maintained, then the engine
must be stopped to prevent damage.
If the oil pressure is too low with associated high oil temperature, then this is probably due to the temperature
causing the viscosity of the oil to drop below the point where the oil pressure relief valve can cope.
If the engine is fitted with oil cooler shutters, then they can be opened in an attempt to lower the
oil temperature. Alternatively, if oil cooler shutters are not an option open to you, then consider either reducing engine power,
or lowering the nose to increase airspeed, or even making the mixture richer.
Any or all of these actions can be used to cool the engine and thus cool the oil, bringing its
viscosity back up to the point where the system can cope.
Once again though, if the minimum oil pressure cannot be maintained, then the engine must be stopped to prevent damage.
Small fluctuations of the oil pressure gauge needle, either side of the desired pressure, may be indicative of a sticking
pressure relief valve.
This would cause the pressure to first build up beyond the optimum level, then subsequently, when the pressure reached a
point high enough to overcome the stiction of the valve, it would open, dropping the pressure momentarily below the correct
level. Then the cycle would start again.
Large fluctuations of the oil pressure gauge needle, on the other hand, will probably be caused by insufficient oil in
the system. The engine may have been leaking oil, and eventually the level will drop to a point where the
pressure pump cannot sustain its output. Then the pressure will drop momentarily until the pump can get some more oil,
when the pressure will rise again, if only for a moment.
Again, if the minimum oil pressure cannot be maintained, then the engine must be stopped to prevent damage.
If the oil pressure falls suddenly to zero, then perhaps the reasons may be a bit academic. After all, if
the minimum oil pressure cannot be maintained, then the engine must be stopped to prevent damage, and there's probably not
a lot you can do about the problem.
Nevertheless, the probable cause is either failure of the pressure pump, or a total loss of oil.
All aircraft piston engines are fitted with dual ignition. That is, they have two electrically independent ignition systems. Each cylinder
therefore has two spark plugs, and each spark plug is fed by a separate magneto.
Dual ignition systems serve two purposes. Firstly, the risk of engine failure due to faulty ignition is reduced, and secondly,
by igniting the mixture within the cylinder at two points, the combustion time is reduced, thus reducing the chance of
detonation.
Magnetos are self-contained engine-driven electrical generators which produce high voltage sparks. Within the magneto, a rotary switch called
a distributor directs the sparks to the plugs in the correct firing sequence.
The magneto combines two basic principles, that of the permanent magnet generator, and that of the step-up transformer.
This diagram illustrates the first basic principle, that of the permanent magnet generator. The only modification to the standard machine
is the addition of a contact breaker which collapses the magnetic field in the primary circuit more quickly than would
normally happen in a generator of this type.
The second basic principle, that of the step-up transformer, is utilised in the secondary circuit. The secondary windings are
a step-up transformer which consists of thousands of turns of very thin wire. A very high voltage is induced
into the secondary windings by the rapid collapse of the magnetic lines of force.
The faster the magnetic lines of force collapse, the greater the voltage induced in the secondary windings.
To ensure that the primary field collapses as fast as possible, and also to prevent the contact breaker points being
eroded by the sparks which would occur when the points opened, a capacitor is placed in parallel with them.
Once the high voltage sparks have been produced, they are passed to the spark plugs in the correct firing sequence.
This is done by the distributor. The distributor rotor arm rotates at half the engine speed.
To give control of the magneto, an ignition switch is fitted in its primary circuit.
Notice that when the ignition switch is placed off, both ends of the primary coil are now earthed, which makes
it impossible for current to be generated.
The secondary circuit is not now cut by the magnetic lines of force which would otherwise enable it to generate
the sparks, so the ignition system is dead.
For illustration purposes, it's supposed that the system shown keeps running in the ignition off period because the propeller is
windmilling.
If, on the other hand, the aircraft was stationary and the ignition switch was left off, the engine would of
course stop.
During start the engine is cranked at about 120 revolutions per minute. At this speed the magneto is not capable
of producing sparks with sufficient energy to ignite the mixture.
This makes it necessary to employ various methods of spark augmentation, of which there are three.
They are the high tension booster coil, the low tension booster coil, and the most common in small aircraft engines,
the impulse coupling.
The high tension booster coil supplies a stream of high voltage impulses to a trailing brush on the distributor rotor.
This ensures that a fat, retarded spark is generated in the cylinder.
The high tension booster coil is only switched on for engine start.
The low tension booster coil supplies the primary coil of the magneto with a low voltage during engine start.
Once the engine has started, the low tension booster coil is switched off.
The impulse coupling is a mechanical device which uses a spring to temporarily increase the speed of rotation of the
magneto. This gives a fat, retarded spark during the starting cycle.
This diagram shows a cutaway view of the impulse coupling within the magneto.
The cam is driven by the gears from the crankshaft at half the crankshaft speed, and drives the hub through
the spring.
The permanent magnet of the magneto is attached to the hub, which also carries the pawls.
Initially the cam rotates turning the spring, which carries the hub round until the pawls come into contact with the
stops.
The cam continues to rotate, tensioning the spring until the leading edges of the cams contact the pawls, pushing them
out of contact with the stops.
The hub, under the impulse of the spring, rotates very quickly round, carrying its attached magnet.
Rotating the magnet this quickly enables it to generate the high voltage spark required to ignite the mixture.
The spark is also retarded, which ensures that the engine is driven in the correct direction.
Once the engine has started, the pawls are rotating at such a speed that centrifugal force holds them out at
an angle which prevents them contacting the stops.
Thus the spring has no further effect.
Although air and fuel vapour will burn when mixed in proportions ranging from between 8 to 1, which is rich,
to 20 to 1, which is weak, complete combustion only occurs with an air-fuel ratio of 15 to 1
by weight.
This is called the chemically correct ratio.
At this ratio, all of the oxygen in the air combines with all of the hydrogen and carbon in the
fuel.
The chemically correct mixture does not, however, give the best results, because the temperature of combustion is so high that
power can be lost through detonation.
Detonation is a phenomenon usually associated with weak mixtures and high cylinder head temperatures.
It's unstable combustion, or more simply, an explosion, which occurs after the normal ignition point.
In normal conditions the flame travels smoothly through the mixture as the advancing flame front heats the gases immediately ahead
of it, so that they in turn burn.
Progressively more and more heat is concentrated in the flame front, which is brought to bear on the remaining unburnt
portion of the mixture termed the end gas, and its temperature is raised.
In addition to this, the end gas is subjected to increasing pressure, partly from the fact that the piston is
reducing the remaining space available, and partly because the burnt gases themselves are expanding.
Ultimately, there can be sufficient pressure and heat available to bring all the end gas to the point of combustion
at the same instant, and it will explode.
The flame rate increases from its normal 60 to 80 feet per second to 1,000 feet per second, with
a degree of violence that depends on the amount of end gas that remains.
Because of the ringing nature of the noise produced by the explosion, it's sometimes called pinking, and it can be
extremely harmful to the engine, particularly the piston crown, which may eventually burn away with the inevitable consequences.
Reducing the temperature of the mixture, for instance by deselecting carburetor heat, or reducing cylinder head temperature by selecting a
rich mixture, will both assist in preventing detonation.
With a supercharged or turbocharged engine, reducing the boost pressure and increasing propeller RPM will also assist in preventing detonation.
Pre-ignition is often confused with detonation, but it is vastly different.
Whereas detonation occurs after the normal ignition point, pre-ignition, as inferred by its very name, occurs before the spark
occurs at the plug.
This is usually caused by a local hot spot in the combustion chamber, either a piece of incandescent carbon or
perhaps very hot spark plug points.
Pre-ignition will cause rough running and an inevitable loss of power.
Although the chemically correct mixture strength would theoretically produce the highest temperature and therefore power, in practice mixing and distribution
are less than perfect, and this results in some parts of the mixture within the inlet manifold system being richer
and others being weaker than the optimum strength.
This variation may exist between one cylinder and another.
A slightly rich mixture doesn't have much effect on power since all the oxygen is still consumed, and the excess
of fuel simply serves to slightly reduce the effective volumetric efficiency. In fact its cooling effect can be to some
extent beneficial.
Weak mixtures however rapidly reduce power since some of the inspired oxygen is not being utilised, and this power reduction
is much greater than that resulting from slight richness.
It is therefore quite common to run engines, when maximum power rather than best fuel economy is the objective, at
somewhat richer than chemically correct mixtures, for example about 12.5 to 1, to ensure that no cylinder is left
running at severely reduced power from its mixture being unduly weak.
However, mixture requirement is ultimately dependent upon engine speed and power output. This diagram shows typical air-fuel mixture curves.
Notice that a rich mixture is required for starting and slow running. This is because fuel will only burn when
it has vaporised and is mixed with air. When starting the engine is cold, and there is little heat to
assist the vaporising process. Therefore only the lightest fractions of the fuel will vaporise.
To make sure that there is sufficient fuel vapor in the cylinders to support combustion, a rich mixture is therefore
required.
When full power is selected for take-off, the mixture must be enriched to about 10 to 1. Apart from
the cooling effect, the excess fuel is wasted, for there is insufficient oxygen available for it to burn completely.
The higher power results from a greater weight of charge induced in a given time, and not because of mixture
enrichment.
In practice, excess fuel vapor is not scavenged as vapor. The oxygen is shared out to some extent, so that
carbon monoxide is produced during combustion, as well as carbon dioxide.
With very rich mixtures, some of the carbon fails to combine with oxygen at all, and is exhausted as black
smoke.
When higher power is required for climbing, the mixture is enriched to about 11 to 1. The extra fuel in
vaporising cools the mixture and reduces the tendency to detonate.
The cruise portion of the diagram shows that it is really the only time the pilot has any choice in
what mixtures are set.
If time is no object, then during cruising flight only moderate power is required from the engine. Thus an economy
cruise setting can be used. Fuel consumption will decrease accordingly.
On the other hand, if speed is of the essence, then performance cruise power can be selected. The fuel consumption
will of course rise, in some cases by as much as 20%.
Because the density of the air decreases with increase of altitude, as the aircraft altitude increases, the weight of air
drawn into the cylinder decreases.
For a given intake velocity, the pressure drop in the carburetor Venturi will decrease as ambient density decreases.
However, the fuel flow due to this pressure drop will not decrease by the same proportion, and so the mixture
will become richer.
If we're not to use too much fuel when flying at altitude, then it becomes imperative that we have some
method of defining and regulating the mixture strength.
There are two approved methods of setting the desired mixture strength. The first method presumes some knowledge of the effects
that mixture change will have on the engine.
Bear in mind that as the mixture control is moved from fully rich to a weaker setting, the air-fuel
ratio approaches the chemically correct value of approximately 15 to 1.
At this ratio, all of the air and fuel are consumed, and the heat released by combustion is at its
maximum. More heat means more power.
With a fixed-pitch propeller, the RPM will rise, and airspeed will increase as more power is produced.
If the mixture is weakened even more, the RPM will drop. Pushing the mixture control back to where the chemically
correct mixture was found will bring the engine speed back up,
but this cannot be maintained without risking the generation of detonation. To prevent this happening, the mixture control should be
moved to a slightly richer position when the RPM will fall slightly.
The exhaust gas temperature gauge consists of a thermocouple fitted into the exhaust pipe of the hottest cylinder on the
engine.
The thermocouple produces a voltage directly proportional to its temperature. The voltage is indicated on a gauge calibrated to show
exhaust gas temperature.
If the mixture control is moved towards lean, the temperature will peak at the ratio of 15 to 1.
It should be remembered that this ratio should not be used as detonation can occur.
Upon reaching the peak EGT, the mixture control should then be moved towards rich, and the cylinder head temperature will
drop.
The aircraft's flight manual will specify a temperature drop which will give the rich cruise setting.
The simplest solution to the problem of maintaining the correct air-field ratio in response to throttle settings at all
operating altitudes is the float chamber carburetor.
The float chamber carburetor employs two basic principles. They are the U-tube principle and the Venturi principle.
The U-tube principle states that if a tube is bent into the shape of a U and then filled
with liquid, the level in either leg will be the same, provided that the pressure acting on each end of
the tube is the same.
If a pressure difference is created across the U-tube, it will cause the liquid to flow.
In the float chamber carburetor, one leg of the U-tube is opened out to form a small tank, and
a constant level is maintained by a float and valve mechanism regulating the flow of fuel from a fuel pump
or pumps, delivering a supply from the main aircraft tanks.
Bernoulli's theorem states that the total energy in a moving fluid is constant.
The fluid possesses energy because of its pressure, temperature and velocity. If one of these changes, one or both of
the others must also change to maintain the same overall energy.
As the air passes through the restriction of the Venturi, its velocity increases, causing a drop in pressure and temperature.
The pressure drop at the throat of the Venturi is proportional to the mass airflow, and this is used to
make fuel flow from the float chamber by placing one leg of the U-tube in the Venturi.
Engine suction provides a flow of air from the air intake through the Venturi in the carburetor to the induction
manifold. This air speeds up as it passes through the Venturi, and a drop in pressure occurs at this point.
Air intake, or atmospheric air pressure, acts on the fuel in the float chamber, which is connected to a fuel
discharge tube located in the throat of the Venturi.
The difference in pressure between the float chamber and the throat of the Venturi provides the force necessary to discharge
fuel into the air stream.
As airflow through the Venturi increases, so the pressure drop increases, and a higher pressure differential acts on the fuel
to increase its flow in proportion to the airflow.
In this simple carburetor, the airflow to the engine is controlled by a throttle valve, and the fuel flow is
controlled by a metering jet.
Although this simple carburetor contains all the components necessary to provide a suitable air-fuel mixture over a limited range,
modifications will be required if it's to give anything more than this basic service.
To maintain the correct rate of discharge of fuel through the main jet, the pressure in the float chamber and
the air intake must be equal.
Equalised pressure conditions can only be obtained by connecting the float chamber directly to the air intake by a duct,
which is called the pressure balance duct.
As engine speed and airflow through the Venturi increase, the proportion of fuel to air rises as a result of
the different flow characteristics of the two fluids.
This causes the mixture to become richer as the engine RPM increases.
In some ways, this may be looked upon as serendipitous. After all, at high engine power, we need a rich
mixture.
However, this is a bit hit and miss, and it doesn't allow any accuracy in setting the mixture.
To overcome this effect, some carburetors are fitted with a diffuser as shown here.
As engine speed is progressively increased above idling, the fuel level in the diffuser well drops, and progressively uncovers more
air holes.
These holes allow more air into the discharge tube, and by reducing the pressure differential, prevent enrichment of the air
-fuel mixture.
The process of drawing both air and fuel through the discharge tube also has the effect of vaporising the fuel
more readily, particularly at low engine speeds, which assists the complete combustion of the fuel in the cylinders.
At low engine speeds, the volume of air passing into the engine is so small that the depression in the
Venturi may be insufficient to draw fuel through the main jet.
A considerable depression exists above the throttle valve, and this is utilised to affect a second source of fuel supply
for slow running conditions.
A slow running fuel passage with its own jet leads from the float chamber to an outlet at the lip
of the throttle valve, as shown here.
The strong depression at this point gives the necessary pressure difference to create a fuel flow.
The size of the slow running jet is such that it will provide the rich mixture required for slow running
conditions.
An air bleed opening into the choke tube below the throttle valve assists atomisation.
The purpose of the transverse passage drilled through the throttle valve is to evenly distribute the mixture over the area
of the induction manifold.
As the throttle is opened, the depression at the lip of the throttle valve decreases, and the depression in the
Venturi increases to the point where the main jet starts to deliver sufficient fuel, and the flow through the slow
running system slows down.
Carburetors must be carefully tuned in order to obtain a smooth, progressive changeover between the slow running and the main
system to prevent flat spots.
A flat spot is a period of poor response to throttle opening caused by a temporary weak mixture.
It normally makes itself felt as a hesitation during engine acceleration.
A cut off valve is usually incorporated in the slow running passage, and is used when stopping the engine.
When the cut off is operated, the valve moves over to block the passage to the slow running delivery.
The mixture being delivered to the engine becomes progressively weaker until it will not support combustion, and the engine stops.
The idle cut off prevents any possibility of the engine continuing to run erratically due to pre-ignition, and also
prevents fuel condensing in the cylinders, which would tend to wash the oil from the cylinder walls, causing lack of
lubrication when the engine was next started.
The cut off may be a separate control, or it may be incorporated in the mixture control lever.
As we stated earlier, as altitude increases, the weight of air drawn into the cylinder decreases, because the air density
decreases.
For a given intake velocity, the pressure drop in the Venturi will decrease as ambient density decreases.
However, the fuel flow due to the pressure drop doesn't decrease by the same amount, and so the mixture will
become richer.
This progressive richness with increased altitude is unacceptable for economic operation.
With a needle type mixture control, such as shown here, a cockpit lever is connected to a needle valve in
the flow chamber.
Movement of the cockpit lever moves the needle in the jet, and varies fuel flow through the main jet.
The position of the needle therefore controls the mixture strength, and in the fully in position will block fuel flow
to the main jet completely, thus providing a means of stopping the engine.
These jets are the smallest orifices in the whole fuel system.
To prevent any blockage of the jets by dirt or debris, a fuel strainer can be fitted just before them
in the fuel line.
The air bleed mixture control shown here operates by controlling the air pressure in the flow chamber, thus varying the
pressure differential acting on the fuel.
A small air bleed between the float chamber and the Venturi tends to reduce air pressure in the float chamber,
and a valve connected to a cockpit lever controls the flow of air into the float chamber.
When this valve is fully open, as shown here, the air pressure is greatest, and the mixture is fully rich.
As the valve is closed, the air pressure decreases, thus reducing the flow of fuel and weakening the mixture.
In the carburetor shown here, the valve also includes a pipe connection to the engine side of the throttle valve.
When this pipe is connected to the float chamber by moving the cockpit control to the idle cut-off position,
float chamber air pressure is reduced, and fuel ceases to flow, thus stopping the engine.
If the throttle valve is opened quickly, air flow responds almost immediately, and a larger volume of air flows through
the carburetor.
The fuel metering system, however, responds less quickly to the changing conditions, and a temporary weakening of the mixture will
occur, known as a flat spot, or at worst, causing a weak cut, before fuel flow again matches air flow.
This condition is overcome by fitting an accelerator pump, which is linked directly to the throttle, and forces fuel into
the venturi whenever the throttle is opened, as shown here.
In some pumps, a controlled bleed past the pump piston allows the throttle to be opened slowly, without passing fuel
to the engine. In other pumps, an additional delayed action plunger is incorporated to supply an additional quantity of fuel
to the engine, for a few seconds after throttle movement has ceased.
The induction system takes air from a forward-facing inlet, which gives the benefit of a slight positive boost from
the aircraft's forward speed, and passes it to the carburetor, and thus to the inlet manifold.
An air filter prevents debris getting through to the carburetor and the engine.
The problems of engine icing, particularly engines fitted with carburetors, have been known for some years, but still accidents occur
in which induction system icing has been the cause, despite modern fuel metering devices.
Atmospheric conditions, particularly of high humidity, a relative humidity greater than 50%, and temperatures ranging from minus 7 degrees Celsius,
that's 20 degrees Fahrenheit, to as high as plus 33 degrees Celsius, 90 degrees Fahrenheit, may cause icing in the
induction system of all types of piston engines.
This temperature range and humidity occur throughout the year in the areas of the United Kingdom and Europe, and therefore
pilots should be constantly aware of the possibilities of icing, and take the corrective action necessary before such problems arise
and the situation becomes irretrievable.
Basically, there are three forms of icing which affect the induction systems of piston engines which are fitted with carburetors.
They are impact ice, refrigeration ice, and fuel icing.
Impact ice forms on bends in the induction system and also on the intake air filter.
Refrigeration ice, or carburetor icing, forms in float-type carburetors as a result of the low temperatures caused by both
fuel vaporization and low pressure acting on the moisture in the atmosphere.
Fuel icing is caused by moisture in the fuel coming out of suspension and being frozen by the low temperatures
in the carburetor. This ice tends to stick to the inlet manifold around its corners, and this reduces the flow
of the mixture into the engine.
Once an engine stops due to induction icing, it's most unlikely that it may be restarted in time to prevent
an accident, and therefore recognition and correction of the problem is vital.
The indications of induction system icing to the pilot of an aircraft fitted with a carburetor and a fixed-pitch
propeller would be a gradual drop in RPM which may be accompanied by engine rough running and vibration.
If the aircraft is fitted with a constant speed propeller, induction system icing would be indicated by a drop in
manifold pressure or reduction of airspeed in level flight.
When icing is suspected, the carburetor heat control should be selected to fully hot and left in the hot position
for a sufficient length of time to clear the ice. This could take up to one minute, or even longer,
depending on the severity of the situation.
Full carburetor heat must always be used unless the aircraft is equipped with a carburetor air temperature gauge. The carburetor
heat control provides heated air from around the exhaust pipe into the induction system, which will melt the ice which
then passes through the engine as water.
Engine roughness and further power loss may occur as the water passes into the cylinders, and pilots should not be
tempted to return the heat control to off, or cold, thinking that the situation has become worse since applying heat.
Icing is also more likely during long periods of flight at reduced power, such as during a glide descent or
letdown for approach and landing. Because the heat is derived from the engine, during long descents the engine temperatures will
gradually cool, thus reducing the effectiveness of the hot air system.
Where icing conditions exist, select full hot air before reducing power so that benefit is gained from the hot engine
before its temperature starts to reduce.
To help maintain engine temperatures and provide a sufficiently great heat source to melt any ice, it's necessary to increase
power periodically to a cruising setting at intervals of between 500 and 1,000 feet during the descent.
Additionally, this action prevents lead fouling of the spark plugs.
Carburetor icing can occur during taxiing at small throttle settings, or when the engine is at idle RPM.
In these circumstances, ensure that hot air is used before take-off to clear any ice.
But remember to select cold air before opening the throttle to full power, and check that the correct take-off
RPM or manifold pressure is obtained. Under no circumstances should carburetor heat be used during take-off.
When using carburetor heat, there are a number of factors which should be understood.
The application of hot air reduces the power output by approximately 15%, and also creates a richer mixture which may
cause rough running.
The continuous use of carburetor heat should be avoided due to the change of mixture and increase of engine temperatures.
Heat should be used only for a sufficient period of time to restore engine power to its original level.
This will be noted by an increase of RPM or manifold pressure above the original setting when the control is
returned to cold.
Don't use the carburetor heat once you're clear of icing conditions, but do check periodically that ice has not reformed.
The following point should be understood when using the carburetor heat control.
Use of the heat control on the ground should be kept to a minimum, as the air is not filtered
and may feed dust and dirt into the system causing additional wear on pistons and cylinders.
A functional check of the heater control should be made before take-off.
Engine rotational speed should drop by approximately 100 RPM when heat is applied, and return to the selected setting when
turned back to off or cold.
If icing is evident on the ground before take-off, use heat to clear the ice, but return the control
to off before applying take-off power.
Check that normal take-off power is available.
When climbing to a new altitude en route, ensure that carburetor heat is selected off.
Be aware of conditions likely to cause carburetor icing, damp, cloudy, foggy or hazy days, or when flying close to
cloud or in rain or drizzle.
Look out for an unaccountable loss of RPM or manifold pressure.
Make frequent checks for icing by applying heat for a period of between 15 to 30 seconds, noting first the
selected RPM, then the drop of RPM as heat is applied.
Listen to the engine noise and check the outside air temperature.
Should RPM increase while heat is applied, or the RPM return to a higher figure than original when reselected to
cold, then ice is present.
Continue to use carburetor heat while flight in icing conditions continues.
During glide descents or long periods of flight at reduced power, say below 1800 RPM, apply carburetor heat, remembering to
warm the engine for short periods, say every 500 to 1000 feet.
During approach and landing, the carburetor heat selector should remain at cold, except for a glide approach, but if icing
conditions are known to exist or even suspected, full heat should be applied.
Remember though, that the carburetor heat selector must be returned to cold before applying power for either a roller landing
or an overshoot.
Engines fitted with fuel injection do not have the problems of ice forming at the Venturi, but other parts of
the system may accumulate ice with a similar loss of power.
Fuel icing may gather at the bends in the system. Impact icing may form at the impact sensing tubes or
on the intake air filters, particularly when flying in cloud at low temperatures.
The alternate air system fitted to these engines should then be selected and the icing drill followed according to the
aircraft checklist.
In the low pressure continuous flow fuel injection system, which is used on many aircraft piston engines, fuel is sprayed
continuously into the induction pipe as close to the inlet valve as possible.
The advantages claimed for the method are low operating pressure, good fuel distribution, freedom from icing problems and the ability
to use a pump which does not have to be timed to the operating cycle.
In this system the fuel metering valve varies the pressure of fuel according to engine speed.
Mixture strength is varied by a manually operated mixture control valve which adjusts the fuel pressure for altitude or operating
conditions as necessary.
Because of the method of operation of the injector, no special idling arrangements are required and a separate priming system
for engine starting is unnecessary.
The main components in the system are a fuel pump, a fuel control unit and of course the discharge nozzles
themselves, one for each cylinder.
In addition, a fuel pressure gauge is fitted to enable any mixture adjustments to be checked accurately.
The fuel pumps supply more fuel than is required by the engine and a recirculation path is provided.
Two pumps are provided arranged in parallel so that when the mechanical pump is not operating, fuel under positive pressure
from the electrical pump can bypass the mechanical pump, so allow the electrical pump to be used for engine priming,
starting and for emergency situations.
A fuel discharge nozzle is located in each cylinder head, with its outlet directed into the inlet port.
Nozzles are fitted to individual engines as a set, each nozzle in a set being calibrated for that set only.
The fuel control unit is mounted on the intake manifold. It contains three control elements, the throttle valve, the metering
fuel valve and the mixture control valve.
The air throttle assembly contains the air throttle valve, which is connected to the pilot's throttle lever and controls airflow
into the engine.
Notice that the intake manifold has no venturi or other restrictions to airflow.
The throttle valve is connected to the pilot's throttle lever and controls airflow to the engine.
The fuel metering valve is connected to the air throttle and controls fuel flow to the fuel manifold valve according
to the position of the air throttle.
Thus fuel flow is proportioned to airflow and provides the correct air-fuel ratio.
The mixture control valve is connected to the pilot's mixture control lever and bleeds off fuel pressure applied to the
throttle metering valve.
Thus the air-fuel ratio can be varied from the basic setting of the throttle metering valve as required by
operating conditions.
Fuel specifications are the means by which producers and users of aviation fuel identify and control the properties necessary for
satisfactory performance.
All aviation fuels then have a specification number, for instance Jet A1, the fuel normally used in jet engines, has
a specification number DERD.
Other jet fuels have different specification numbers.
Gasoline is the fuel normally used in aviation piston engines.
Gasoline has a specification number DERD2485.
Contained within that blanket specification number are grades of gasoline such as Avgas 100, Avgas 100LL and Avgas 80.
Many tests are carried out on the fuel when it's manufactured.
One of those tests is used to determine the octane rating.
To do this the particular blender fuel that we are interested in is compared with two reference fuels under standardised
conditions in a special test engine.
The two reference fuels are called isooctane and normalheptane.
Isooctane has very good combustion characteristics and shows little tendency to detonate when mixed with air and ignited at high
temperatures.
It's given a rating of 100.
Normalheptane on the other hand detonates very readily and has a rating of 0.
The particular fuel that we are interested in is compared with a blend of the two reference fuels.
This is done by first running our fuel in the special test engine and then trying to produce the same
degree of detonation in the engine while using a blend of the two reference fuels.
If the blend of the two reference fuels which gives the same detonation characteristics is say 95% isooctane and
5% normalheptane, then our fuel under test would be given an octane rating of 95.
The octane rating is considered to be a measure of how well the fuel resists detonation or its anti-knock
value.
Originally tests were based on an air-fuel ratio which gave maximum detonation, but this condition is not truly representative
of the working range of the engine.
Maximum detonation occurs with economical mixtures which are normally used for cruising, but for take-off and climbing rich mixtures
are used.
It's important to know how the fuel will behave under these varying mixture strengths, and so aviation fuel has two
ratings.
This is sometimes referred to as the performance number or performance index.
As an example, ABGAS 100 is a 100-octane fuel with a performance number of 100-130.
The lower figure is the weak mixture detonation point, and the higher figure the rich mixture detonation point.
It follows that if an engine is designed to use a certain grade of fuel, then a lower grade should
never be used, as this would cause detonation.
If at any time the correct octane rating is not available, then a higher octane rating must be used.
To increase the octane rating of a fuel, tetraethyl lead, or TEL, used to be added.
For instance, two grams of lead were added to each gallon of fuel to take its octane rating to 115
,145.
The action of tetraethyl lead is to reduce the formation of peroxides, which act as fulminates to explode the end
gas,
but it has to be used with care, as during combustion, lead oxide is formed.
Lead oxide is not volatile at these temperatures, and it has a corrosive effect on the exhaust valve, its seat,
and the spark plug electrodes.
To prevent this corrosion, it's necessary to add ethylene dibromide to the fuel to change the reaction during combustion to
allow lead bromide to form.
Lead bromide is volatile, and thus is easily ejected, as a halide, with the exhaust gases.
Lead in the atmosphere is absorbed into the bloodstream, especially organic forms, such as lead bromide, which are fat soluble.
Because they're fat soluble, they therefore have a particular tendency to distribute themselves into the brain, and will cause damage
there.
To make it easy to identify the fuels, they are coloured differently from each other.
For instance, Avgas 100LL is coloured blue, while Avgas 100 is coloured green.
Avter, which is aviation turbine fuel, is either clear or is a straw colour.
To assist in differentiating between the fuels when refuelling your aircraft, all refuelling equipment is marked in such a way
as to make it easy to identify which fuel may be dispensed from a particular device.
The colours of the labels on the equipment and the pipelines are shown here.
Some aviation authorities do allow the use of automobile gasoline, or MOGAS, in some aircraft.
However, within the United Kingdom, the rules governing the use of MOGAS are laid down in Airworthiness Notices numbers 98
and 98A.
Further information on the use of MOGAS can be found in the CAA Safety Sense Leaflet number 4A.
Great caution should be taken when using MOGAS. It has a much higher volatility than Avgas, and consequently will evaporate
much more quickly.
This means that the possibility of vapour locks and carburettorising occurring is much greater than would otherwise be if Avgas
was being used.
The most common contaminant in fuel is water. Water is always present in the fuel in varying amounts despite the
manufacturer's stringent quality control and the preventative measures taken during storage and transfer.
However, further message can be taken to minimise water accretion once the fuel has been transferred to the aircraft tanks.
Once the fuel is in the aircraft fuel tanks, the main source of water contamination comes from the atmosphere that
remains within the tank.
If the tanks are topped up to full, then the atmosphere is excluded, together with the moisture it contains, thus
minimising the likelihood that the fuel will be contaminated.
If the fuel can be allowed to settle after replenishment, then the water droplets, being heavier than the fuel, will
fall to the bottom of the tank and can then be drained off through the water drain valve.
Each fuel tank is equipped with a quick drain. This is located at the lowest part of the tank, sometimes
within a sump, which will collect any water as it settles.
These wing tank quick drains should be used first to drain the sumps.
Sufficient fuel should be allowed to flow to ensure removal of contaminants.
This fuel should be collected in a suitable container, examined for contaminant, and then discarded.
After draining any amount of fuel, always ensure that no fire hazard exists before starting the engine.
After using each drain, always ensure that it has been closed completely and that it is not leaking.
While you're in the vicinity of the fuel tank drains, you might as well check the fuel tank vents.
Each tank is fitted with a forward-facing vent pipe, which allows atmospheric pressure to be maintained inside.
They are usually fitted in the lower surface of the wing.
Make sure they're not damaged or that they're not blocked up in any way, otherwise a depression will form in
the tank and will prevent the fuel getting to the engine from that tank.
The fuel strainer is also equipped with a quick drain facility.
The fuel strainer itself is located at the lowest point of the fuel system, usually somewhere in the engine compartment.
The fuel strainer should be drained once for each of the tanks which can be selected on the aircraft's fuel
selector valve.
Once again, sufficient fuel should be allowed to flow to ensure complete removal of any water.
A simple light aircraft fuel system is shown here.
The fuel is contained in two wing tanks, each of which we have seen as a quick drain fitted at
its lowest point.
Fuel quantity is gauged by a simple float-type sensor and is shown in the cockpit on two separate dials.
Float-type sensors are fairly unreliable, except either on the ground or in prolonged straight and double flight.
So before flight always visually check the fuel contents by removing the tank filler caps and looking into the tanks
to make your own assessment of the quantity.
The fuel passes to the engine system via a tank selector valve.
This allows the pilot to draw fuel from each tank in turn and thus keep the aircraft balanced laterally.
Next in the line is the fuel strainer, at the lowest point in the system. Remember, the fuel strainer should
be drained twice via its quick drain, once with a fuel selector valve on each tank.
The electrical fuel pump is next in line. It's sometimes called the booster pump or the auxiliary pump.
It's fitted to the system in case of failure of the mechanical fuel pump.
The pilot's operating handbook for most light aircraft advises that the electrical fuel pump should be switched on for take
-off, flight below 1000 feet and for landing.
It's also advisable to select this pump on when using the tank selector valve to change tanks.
Notice that there is a fuel pressure gauge fitted to the instrument shown. This can be used to check the
output of the electrical pump when the engine is not running.
Downstream of the electrical pump is the engine driven mechanical pump. The fuel pressure gauge will sense the output pressure
of the mechanical pump whenever the electrical pump is switched off.
High wing aircraft with tanks only in the wings may not require a fuel pump, leaving gravity as the prime
mover of the fuel to the engine system.
The final item in the system is the carburetor. This meters the fuel to the engine through the authority of
the throttle and the mixture control.
The priming pump is a hand operated device which is used to pump fuel to the inlet valve ports of
the cylinders prior to engine start.
The priming pump takes its fuel from the top of the fuel strainer. Great care must be taken to ensure
that the priming pump plunger is locked in after use.
If it inadvertently becomes unlocked it may vibrate open. This will create a risk of fuel being sucked from the
top of the fuel strainer into the inlet manifold of the engine, thereby making the mixture extremely rich and perhaps
stopping the engine.
The purpose of a propeller is to convert the power delivered by an engine into propulsive thrust in order to
propel an aircraft.
This thrust is achieved by the acceleration of a comparatively large mass of air rearwards, thereby producing forward thrust.
The acceleration applied is not large when compared with other reaction systems, but this has its advantages, not least of
which is that it is very efficient.
The propeller consists of two or more aerodynamically shaped blades attached to a central hub. This hub is mounted onto
a propeller shaft driven by the engine.
The whole assembly is rotated by the propeller shaft rather like rotating wings.
Like a wing a propeller blade has a root and a tip, a leading and trailing edge, and a cambered
cross section whose chord line passes from the centre of the leading edge radius to the trailing edge.
At the root area where the section of the blade becomes round is the blade shank, while the base of
the blade, where any pitch change mechanism would have to be attached, is called the blade butt.
Having stated that the propeller blade is an aerofoil it would be logical to expect that the definitions relating to
the propeller would be the same as those for the wing. They are as follows.
The chord line is a straight line joining the centres of curvature of the leading and trailing edges of an
aerofoil.
The chord is the distance between the leading and trailing edges measured along the chord line.
There are however other definitions which are peculiar to the propeller and we must consider these, for instance.
The blade angle or pitch is the angle between the blade chord and the plane of rotation.
Blade angle decreases from the root to the tip of the blade because the rotational velocity of the blade increases
from root to tip.
For reference purposes the blade angle is measured at a point 75% of the blade length from the root.
Sections near the tip of the propeller are at a greater distance from the propeller shaft and travel through a
greater distance.
Tip speed is therefore greater. The blade angle must be decreased towards the tip to give a constant angle of
attack along the length of the blade.
The blade angle determines the geometric pitch of the propeller. A small blade angle is called fine pitch, a large
blade angle is called coarse pitch.
The geometric pitch is the distance the propeller would travel forward in one complete revolution if it were moving through
the air at the blade angle.
In flight the propeller does not move through the air at the geometric pitch.
The distance it travels forward in each revolution depends on the aircraft's forward speed.
The distance which it actually moves forward in each revolution is called the effective pitch or advance per revolution.
The difference between the geometric pitch and the effective pitch is called propeller slip.
The helix angle is the angle that the actual path of the propeller makes to the plane of rotation.
Although it definitely has connotations with regard to the wing, angle of attack is also pertinent to the propeller.
The angle between the blade chord and the relative airflow is the angle of attack, shown here as alpha.
The angle of attack is the result of the propeller's RPM and the aircraft's true airspeed.
Notice how the angle of attack changes as the prop RPM changes.
A propeller must be able to absorb all the shaft horsepower developed by the engine and it also must operate
with maximum efficiency throughout the required performance envelope of the aircraft.
The critical factor here is tip velocity. If tip velocity is too high the blade tips will approach the local
speed of sound and compressibility effects will decrease thrust and increase rotational drag.
Supersonic tip speed will considerably reduce the efficiency of a propeller and greatly increase the noise it generates.
This imposes a limit on propeller diameter and RPM and the true airspeed at which it can be used.
Another limitation on propeller diameter is the need to maintain adequate ground clearance.
To increase power absorption several characteristics of the propeller can be adjusted.
The usual method is to increase the solidity of the propeller.
Propeller solidity is the ratio of the total frontal area of the blades to the area of the propeller disc.
An increase in solidity can be achieved by firstly increasing the chord of each blade.
This increases the solidity but blade aspect ratio is reduced, making the propeller less efficient.
Secondly, the number of blades can be increased.
Power absorption is thus increased without increasing the tip speed or reducing the aspect ratio.
However, increasing the number of blades beyond a certain number, say 5 or 6, will reduce overall efficiency.
Thrust is generated by accelerating air rearwards.
Making the disc too solid will reduce the mass of air that can be drawn through the propeller and accelerate
it.
To increase the number of blades efficiently, contra-rotating propellers are used.
That is, two sets of blades rotating in opposite directions on the same shaft.
Due to its rotation, a propeller generates yawing, rolling and pitching moments.
These are due to several different causes such as torque reaction, gyroscopic precession, spiral slipstream effect and asymmetric blade effect.
The majority of modern engines are fitted with propellers which rotate clockwise when viewed from the rear, so called right
hand propellers.
If then the propeller rotates clockwise, the equal and opposite reaction, or torque, will give the aircraft an anti-clockwise
rolling moment about the longitudinal axis.
During take-off this will apply a greater download to the left wheel, causing more rolling resistance on the left
wheel making the aircraft want to yaw to the left.
In flight, torque reaction will also make the aircraft want to roll to the left.
Torque reaction will be greatest during high power, low indicated airspeed flight conditions.
Low indicated airspeed will reduce the power of the controls to counter the turning moment due to torque.
Torque reaction can be eliminated by fitting contra-rotating propellers.
Torque from the two propellers, rotating in opposite directions on the same shaft, will cancel each other out.
A rotating propeller has the properties of a gyroscope, rigidity in space and precession.
The characteristic which produces gyroscopic effect is precession.
Gyroscopic precession is a reaction that occurs when a force is applied to the rim of a rotating disk.
When a force is applied to the rim of a propeller, the reaction occurs 90 degrees ahead in the direction
of rotation and in the same direction as the applied force.
As the aircraft is pitched up or down, or yawed left or right, a force is applied to the rim
of the spinning propeller disk.
For example, if an aircraft with a clockwise rotating propeller is pitched nose up, the reaction occurs at 90 degrees
in the direction of rotation.
That is, a right yawing moment.
Gyroscopic effect can be easily determined when the point of application of the imagined forward force on the propeller disk
is considered.
As the propeller rotates, it produces a backward flow of air, or slipstream, which rotates around the aircraft as illustrated
here.
This spiral slipstream causes a change in airflow around the fin.
In the case shown here, due to the clockwise direction of propeller rotation, the spiral slipstream meets the fin at
an angle from the left, producing a sideways force on the fin to the right.
The amount of rotation given to the air will depend on the power setting.
Spiral slipstream effect can be reduced by any or all of the following.
A small fixed tab on the rudder, the engine thrust line being inclined slightly to the right, or offsetting the
fin slightly.
A fixed pitch propeller receives its relative airflow from a direction governed by the aircraft's true airspeed in the direction
of flight and its own RPM in the plane of rotation.
The operating angle of attack will be the angle between the relative airflow and the chord line of the propeller
blade.
This chord line will be set at an angle to the plane of rotation, that is the blade angle or
the propeller pitch angle.
This animation shows that an increase in RPM will increase the angle of attack, whereas an increase in true airspeed
will reduce it.
At high forward speed with low RPM, for instance in a power-off dive, it's possible to reduce the angle
of attack to zero, while at a low true airspeed with a high RPM for instance in a climb, it's
possible to stall the propeller blade.
Both extremes are obviously inefficient and therefore undesirable.
The conclusion that must be drawn is that for a given fixed pitch a propeller will only work efficiently at
one combination of true airspeed and RPM.
All these considerations illustrate why more advanced aircraft have engines which are fitted with variable pitch propellers.
There are several different types of variable pitch propellers.
Here are some which the private pilot may come across.
Adjustable pitch propellers are propellers which can have their pitch adjusted on the ground by mechanically resetting the blades in
the hub.
In flight they act as fixed pitch propellers.
Two pitch propellers are propellers which have a fine and coarse pitch setting that can be selected in flight.
Fine pitch can be selected for take-off, climb and landing, and coarse pitch for cruise.
Two pitch propellers will usually also have a feathered position.
Modern aircraft have propellers which are controlled automatically to vary their pitch, the blade angle, so as to maintain a
selected RPM.
A constant speed variable pitch propeller permits high efficiency to be obtained over a wider range of true airspeeds, giving
improved take-off and climb performance and cruising fuel consumption.
This diagram illustrates a typical set of engine and propeller controls for a small piston engine aircraft with a variable
pitch propeller.
Throttle, propeller and mixture are shown in the take-off, all forward position.
For a given throttle setting, pulling back on the propeller control will decrease RPM by coarsening the pitch of the
blade.
Conversely, pushing forward on the propeller control will increase RPM by fining off the pitch of the blade.
Once RPM has been set, it will be maintained within the operating range of the throttle lever.
This shows conditions during the early stages of take-off roll.
The RPM is set to maximum and the true airspeed is low.
The angle of attack is optimum and maximum available efficiency is obtained.
As the aircraft continues to accelerate the true airspeed will increase, which decreases the angle of attack on the blades.
Less thrust will be generated and less propeller torque.
This gives less resistance for the engine to overcome and RPM would tend to increase.
The constant speed unit or CSU senses the RPM increase and increases pitch to maintain the blade angle of attack
constant.
This diagram shows the conditions at high forward airspeed in level flight.
As the true airspeed increases, the constant speed unit continually increases the blade angle or cautions the pitch to maintain
a constant blade angle of attack.
If a loss of engine torque occurs, the throttle is closed or the engine fails, the propeller will fine off
in an attempt to maintain the set RPM.
The relative airflow will impinge on the front surface of the blade and generate drag and negative propeller torque.
The propeller will now drive the engine. The drag generated by a windmilling propeller is very high.
A single engine aeroplane fitted with a constant speed propeller will not have a feathering capability as such.
However, following engine failure drag can be reduced to a minimum by moving the propeller pitch control to the fully
course position, as shown here.
In a steady glide with no shaft power from the engine, the throttle closed, if the propeller pitch is increased
by pulling back the prop lever, the aircraft lift drag ratio will increase.
This will decrease the rate of descent. The RPM would decrease because of the reduction in negative propeller torque.
In the interests of the safety of the aircraft, people on the ground and that of the pilot, many precautions
have to be taken before and after flight.
We'll begin with the pre-start situation.
The aircraft should be sighted on firm level ground, free from loose stones and litter, and servicing equipment that might
be damaged by the slipstream.
Where practicable, the aircraft should be heading nose into wind, and so positioned that the slipstream is directed away from
other aircrafts and buildings, and that there is room to taxi.
The brakes should be applied, or alternatively the main wheels should be securely chocked. All covers and blanking plugs must
be removed before starting the engine, and all access panels and hatches must be securely closed.
The engine oil content should be checked to ensure that there is sufficient oil for the planned flight.
Some aircraft have the oil dipstick integral with the oil filler cap, so make sure that the filler cap is
secure after contents checking or replenishment.
Look for signs of oil, fuel or hydraulic fluid leaks.
Check that plug leads and priming lines are secure.
Once the check inside the cowlings is finished, make sure that the cowlings are closed and the catches are secure.
The propeller should be checked for damage. Make sure that there are no nicks or cuts on the blade.
Check that the spinner is secure and undamaged.
Treat the propeller with a great deal of respect at all times.
The results of the engine firing just once could be fatal if you are in the way of the propeller
blade, so stay out of its arc and don't touch it or move it unnecessarily.
If you have any assistance to start the aircraft, make sure that those individuals involved are thoroughly conversant with the
procedures used.
Also ensure that they're not wearing loose clothing that may get tangled in the propeller or blown about the airfield
in the slipstream.
Now let's look at engine start. Bear in mind that the advice given here is general.
It doesn't apply to any particular aircraft type.
Initially, prime the engine with the recommended amount of fuel for the prevailing engine temperature.
Being too generous with the priming may cause trouble for you later.
Firstly, it may flood the engine and wet the plugs, which will definitely not help the engine to start.
And secondly, it could also put fuel into the induction system, which may start a carburetor fire if the engine
backfires.
This is potentially a very dangerous situation.
If you are unfortunate enough to have a carburetor fire, keep the engine turning over on the starter motor, put
the mixture control to the idle cutter position and open the throttle.
Turning the engine over will tend to suck the flames into the cylinders where they won't do any harm.
Selecting the idle cut-off will starve the fire of fuel and opening the throttle will make it easier for
the engine to suck the flames into the cylinders.
Continuing with the normal start, set the mixture control to the rich position and set the throttle just open.
In the interest of safety, it's best at this point to check that the area around the propeller is free.
Making sure that all in the vicinity can hear you, shout Clear Prop!
Now turn the magneto starter switch through to the start position.
Most aircraft starter motors have a limitation on how long they can remain energised.
Make sure you know that limitation and be sure that you don't exceed it.
As soon as the engine is running, release the start switch and immediately check that the starter warning light is
out.
If the starter warning light is still illuminated, stop the engine immediately.
Now check the engine oil pressure gauge.
If the oil pressure doesn't rise within 30 seconds, stop the engine immediately.
The engine should now be allowed to warm up until minimum temperatures and pressures are obtained before the throttle can
be opened.
This warming up is done at the recommended RPM stated in the operator's handbook.
This is a convenient time to check that the aircraft systems are functioning satisfactorily.
For example, check that the fuel pressure is within limits and that the suction system is providing the correct amount
of vacuum.
With the throttle still set at the prescribed slow running RPM, the dead cut check can be carried out.
This check is endeavouring to ascertain whether the two magnetos are both working under your control.
Select each magneto in turn.
Check in each case that there is a drop in RPM but that the engine does not stop.
If the engine does appear to fail when either magneto is selected, allow it to stop completely.
Don't be tempted to select the magneto switch either to the other magneto or back to the both position.
This could possibly cause mechanical damage to the engine.
Now check that the engine pressures and temperatures are within the green, or at the very least, the yellow sectors
of the gauges.
Change the fuel tank selector to the other tank and set the throttle to the correct RPM for the magneto
check, usually around 2000 RPM.
Check the carburetor heat now.
You should note a drop in engine speed as it is selected on because the engine is receiving hot, less
dense air, so the mixture has been made richer by your selection.
The RPM lost should be regained after carburetor heat has been deselected and once more cold air is being fed
into the carburetor.
Now on to the magneto check proper.
The magneto check is carried out to make sure that both magnetos are functioning correctly.
As each magneto is switched off in turn, check for a drop in RPM.
This drop must be within the limits laid down by the manufacturers, and the mathematical difference between the RPM drop
of each magneto must also be within a tolerance figure which is given once again by the manufacturers.
The fall in engine speed is due to the fact that as you switch off a magneto you're essentially switching
off a plug in each cylinder, thus increasing the time it takes for the mixture to be fully consumed within
the cylinder.
If no drop in RPM occurs, then this may be indicative of the fact that the deselected magneto has not
been deselected or it may be that it wasn't working in the first place.
However, that should have been noticed during the dead cut check.
An excessive RPM drop when a magneto is switched off may be indicative of any of several problems.
For instance, the plugs of the remaining magneto may be defective or just fouled.
If fouled plugs are suspected, then it may be worth leaning the mixture off for about 10 to 20 seconds
to clear them, and then resetting the mixture to rich, preparatory to trying the check again.
After completing the magneto check, check the aircraft systems once again.
Note the engine oil temperature and pressure gauges, and the low oil pressure warning light.
Ensure also that the fuel pressure is within tolerances, and that the suction vacuum is sufficient.
After checking that the alternator is still on-line, and that the ammeter reading is normal for the load selected,
close the throttle and note that the engine idles at about 600 RPM, and then set it to the recommended
idle speed, usually about 1200 RPM.
Setting the throttle at this level prevents the plugs becoming fouled during prolonged idling.
Get used to operating the throttle lever smoothly.
Jerky or rapid movement of the throttle lever will not make the engine respond any more quickly.
In fact, it may cause the engine to hesitate or stutter because of an over-weak mixture.
Engines have been known to stop because of a weak cut caused by a rapidly open throttle.
Count slowly up to three while opening the throttle from idle to full power.
You'll find that the engine is accelerated at about the same rate as the throttle lever is moving, which is
about as fast as it can respond.
The same practice should be followed when reducing power, especially when throttling back from take-off power settings.
Smooth operation of the engine controls will allow time for mixture strength and charge quantity to change in-line with
the engine's requirements.
The exhaust gas temperature gauge and the cylinder head temperature gauge are often confused with each other.
However, they do perform different functions.
The exhaust gas temperature probe is installed about 4 inches from the cylinder head on the exhaust system.
Although the exhaust gas temperature gauge can help in troubleshooting a problem on the engine, it's primarily a fuel management
instrument.
On the other hand, the cylinder head temperature gauge is an engine instrument designed to protect the engine against its
enemy – excessive heat.
Most general aviation aircraft engines take the cylinder head temperature off the hottest single cylinder of the four, six or
eight cylinders.
The hottest cylinder is determined by extensive flight tests carried out by the engine manufacturers.
Minimum in-flight cylinder head temperature should be about 150 degrees Fahrenheit, and maximum in most direct-drive normally aspirated
engines is 500 degrees Fahrenheit.
Some of the higher-powered, more complex engines have a maximum limit of 475 degrees Fahrenheit.
Although these figures are minimum and maximum limits, the pilot should operate his engine at more reasonable temperatures in order
to enable it to achieve its expected overall life.
In practice, it's been found that engines have benefited greatly during continuous operation by keeping cylinder head temperatures below 400
degrees Fahrenheit in order to achieve the best life and wear.
In general, it would be normal during all-year operations in climb and cruise to control the cylinder head temperatures
in the range of 350 degrees to 440 degrees Fahrenheit.
The way the mixture control works has been described previously in the carburetion lesson.
We'll now describe its operation during the take-off and landing phase, the climb and cruise, and also its use
as an idle cut-off.
We know from the carburetion lesson that detonation is most likely to occur under conditions of high-charge temperatures and
pressures.
The take-off phase of a flight is exactly when these conditions will occur, and also, as it could be
likely that we may have to go around from the landing phase, once again generating the same conditions as on
take-off, we must select a fully rich mixture in both circumstances.
The addition of extra fuel into the cylinders reduces the volumetric efficiency of the engine, which somewhat reduces the potential
of the engine to produce power, but on the other hand, the cooling effect that the evaporation of the fuel
has on the charge prevents detonation and therefore allows the engine to deliver its maximum rated output.
We also found from the carburetion lesson that because of the air's decreasing density with increasing altitude, our mixture gets
richer as the aircraft climbs, because the weight of fuel being fed into the cylinders doesn't reduce by the same
proportion.
As long as the climb doesn't continue above say 5,000 feet, this isn't really a problem. The rich mixture
does the same job it did during the take-off phase. It stops detonation by cooling the charge temperature, so
the mixture control can be left at fully rich.
During the cruise phase of flight, the mixture can be leaned off to give better fuel economy. If your aircraft
has an exhaust gas temperature gauge, then this can be used to give an indication of the correct mixture setting.
If the mixture control is moved towards lean, the temperature will peak at the ratio of 15 to 1. It
should be remembered that this ratio must not be used, as detonation will occur.
On reaching the peak exhaust gas temperature, the mixture control should then be moved towards rich, and the cylinder head
temperature will drop. The aircraft's flight manual will specify a temperature drop, which will give the rich cruise setting.
If an exhaust gas temperature gauge is not available, then this alternative method of setting the correct mixture strength for
the cruise may be utilised. Moving the mixture control from fully rich to a weaker setting brings the air-fuel
ratio closer to the chemically correct value of approximately 15 to 1.
Remember, at this ratio all of the air and fuel are consumed, and the heat released by combustion is at
its maximum. More heat means more power, so with a fixed-pitch propeller the RPM will rise, and airspeed will
increase as more power is produced.
Weakening the mixture from this point will make the RPM drop. Pushing the mixture control back towards rich, to where
the chemically correct mixture was found, will bring the engine speed back up, but this cannot be maintained without risking
the generation of detonation.
To prevent this happening, the mixture control should be moved back to a position richer than that where the peak
RPM was found. When the RPM will at first rise as it passes through the 15 to 1 setting, and
then fall slightly, this position can be assumed to be the optimum for a rich cruise.
The normal method of shutting down an aircraft engine is by cutting off the fuel supply. This has several advantages
over the method used in automobile engines, where the ignition is switched off instead. Cutting off the fuel ensures that
the engine cannot run on, a phenomenon which can occur if the engine has any incandescent carbon deposits in the
combustion chamber.
It also means that no fuel is allowed to be sucked into the cylinders during those last few revolutions before
the engine stops. Any fuel brought into the cylinder after combustion ceases will tend to wash the lubricating oil from
the cylinder walls, and as a consequence, the next time it starts, the engine suffers a tremendous amount of wear.
In the carburetion lesson the idle cut-off section of the carburetor was described in some detail.
The idle cut-off may be a separate control, or it may be incorporated into the mixture control lever. By
pulling the mixture control lever to the ICO position, the fuel flow between the float chamber and the venturi is
cut off.
Most light aircraft use a direct current, or DC supply, to drive the electrical services. The term direct current implies
that the electrons which make up the flow of current are travelling in one direction only. Convention states that the
electrons flow from a positive source to a negative.
Alternating current, on the other hand, has the electrons in the circuit moving backwards and forwards about a mean point.
The number of electrons flowing past any one point in a conductor is measured in amperes, or amps for short,
on an ammeter. Imagine the current of electrons flowing through a wire as being the same as water flowing in
a pipeline. This is called the water analogy, and for people not too well acquainted with electricity, the water analogy
can come in very handy in almost all situations when dealing with simple direct current circuits.
So, as with any simple plumbing system, for the water to flow through the pipe requires that we have a
pressure. This pressure is usually obtained from a water tank, which is placed above the level of the whole water
system, and thus gravity helps us out.
In an electrical circuit, the pressure required to move electrons along a wire is termed the electromotive force, or the
voltage, which is measured in volts, on a voltmeter. We can obtain this electromotive force, or voltage, from any of
several sources.
In a light aircraft, the main source of supply is the generator, or alternator, and the battery serves as a
backup in case of alternator failure. The battery can run the aircraft's essential services for about 30 minutes. The battery
also supplies power for engine starting.
Resistance is the obstruction in a circuit which opposes current flow. Using the water analogy, you could imagine that a
kink in a hose pipe would restrict the water flow. Likewise, a resistor placed in a circuit would lower the
current flowing in that circuit.
The supply is fed to the electrical components of the system via buzz bars, which are merely collection and distribution
points. The buzz bars are usually solid copper bars which are drilled and tapped so that supply and distribution cables
can be attached to them.
Most light aircraft which utilise metal construction are single pole, or earth return, electrical systems. This means that the individual
components are supplied via the buzz bars and cables, and to complete the circuit the return current flows back through
the metal of the airframe.
Aircraft which are made of non-conductive materials require a double pole, or two wire system. This means that as
well as having a cable taking the current flow to each individual component, another cable is required to complete the
electrical circuit back to the negative side of the generator or alternator.
This is probably a good time to explain the use of the two terms alternator and generator. Simply put, a
generator produces direct current, while an alternator produces alternating current internally, using a diode rectifier to turn the alternating current
into direct current, which is then fed to the aircraft circuit.
Alternators are used much more in all applications where a dependable power supply is required. Whereas a generator will probably
require the engine to run at approximately half speed before it will deliver its full output, an alternator will give
almost full power at engine idling speed.
A battery is made up of a number of cells which convert chemical energy into electrical energy by a transfer
of electrons from one material to another.
There are two types of cell, the primary and the secondary. The primary cell is the type normally used in
hand-held equipment, like torches and transistor radios etc.
It consists of two electrodes immersed in a chemical called an electrolyte. The electrolyte encourages electron transfer between the electrodes
until the electron imbalance between them exists as a potential difference or voltage.
A voltage of approximately 1.5 exists between the electrodes of a new primary cell.
For convenience, one of the electrodes is used as the container for the electrolyte, and the other electrode is placed
within it, as shown here.
When the two electrodes, now called the positive and the negative, are connected to an external circuit, electrons flow through
the circuit.
This external transfer of electrons is matched by an internal transfer from the positive to the negative electrode.
As this circulation of electrons continues, the negative electron slowly dissolves into the electrolyte until it is eventually eaten away.
At this point the cell is dead.
Once discharged, primary cells cannot be recharged.
Secondary cells work on the same principle as primary cells, but the chemical energy in the cell can be restored
when the cell has been discharged by passing a charging current through the cell in the reverse direction to that
of the discharge current.
In this way the secondary cell can be charged and discharged many times over a long period of time.
During charging, electrical energy is converted into chemical energy which is retained within the battery until it is once again
discharged.
The lead-acid battery is most commonly used in light aircraft.
The electrodes in a lead-acid battery are called the plates.
There is a positive plate, which is made of lead peroxide, and a negative plate, which is made of spongy
lead.
Both of the plates are immersed in an electrolyte of dilute sulfuric acid.
The state of charge of a lead-acid cell can be determined by measuring the specific gravity of the electrolyte
solution with a hydrometer.
The electrolyte of a fully charged cell will have a specific gravity of approximately 1.3, whereas the electrolyte of
a discharged cell will have a much more dilute acid with a specific gravity somewhere in the region of 1
.17.
The voltage of a lead-acid battery cell when it is offload or not passing current is 2.2 volts.
Apply a load to the terminals of the cell and the output voltage will fall to just 2 volts.
The capacity of a cell is a measure of how much current it can provide over a certain period of
time.
The capacity of a cell is determined by the area of the electrodes or plates as they are called in
a secondary cell.
Capacity is measured in ampere hours.
A cell with a capacity of 80 ampere hours should provide a current of 8 amperes for 10 hours, or
alternatively 80 amperes for 1 hour.
However, because the ability of a battery to retain its output voltage diminishes as the discharge rate increases, capacity is
normally tested at the 10 hour rate.
It is not normal to use cells in isolation in aircraft batteries. Usually they are placed in series or in
parallel with others of the same voltage and capacity.
This cut-away diagram shows six cells in series with each other.
This circuit diagram shows cells connected in series. Notice how the voltage of the circuit increases as more cells are
added.
Note also, however, that the capacity of the system remains the same as the capacity of one cell.
This diagram shows cells joined together in parallel.
Now the output voltage of the circuit is that of just one cell.
But notice that the total capacity is the sum of the individual cell capacities.
To charge a battery it must be connected to an electrical supply which has a slightly higher voltage than that
of the battery itself.
When the battery is being charged, great care must be taken regarding the rate of charge. If the rate of
charge is too high, then gassing will take place.
Gassing is the formation of hydrogen gas on the plates.
The gas must be allowed time to escape through the vent cap of the battery.
If too much gas is produced in too short a time, then there is a danger that a build-up
of gas pressure could cause the battery to explode, with the subsequent release of the sulphuric acid within.
Under normal circumstances, the output voltage of the alternator or generator charging a 12 volt battery on an aircraft would
be maintained constant at 14 volts.
Whereas the alternator of an aircraft with a 24 volt system would be maintained constant at 28 volts.
After starting an engine using the aircraft battery, the generator or alternator is used to recharge it.
The fact that the battery is being recharged is indicated on the system ammeter as a high load.
Initially this will be quite high, but should quickly reduce as the battery is recharged.
A word of caution here. If the load increases or remains high, then this could be an indication of a
faulty battery.
Remember, a high charge rate could result in a battery overheating and suffering damage.
Voltmeters have a high internal resistance. They are connected in parallel with the circuit to measure the voltage between two
points.
Ammeters have a low internal resistance and are placed in series with the circuit to measure the current through the
load.
Two types are used in light aircraft. The first has the zero position on the dial over to the left.
It indicates the actual load on the alternator. Because of this it is commonly called a loadmeter.
The other type has the zero position in the centre of the dial and its needle indicates the current flow
into and out of the battery.
Abnormal conditions may arise in an electrical circuit for a variety of reasons.
For instance, a breakdown of the insulation on one of the cables carrying the supply to a component will cause
a dramatic rise in the current flowing in that cable if it comes into contact with the metal of the
airframe.
There are several serious problems inherent in this situation.
Firstly, and probably least serious, is the fact that the component being fed by this particular cable will cease to
operate.
Secondly, if the circuit was not protected, the current which would flow through this short circuit, for that's what this
malfunction is called, would be so great that it could cause the power generation circuit, the alternator or generator and
the battery, to fail.
This would leave the aircraft without electrical power.
Thirdly, and potentially the most serious, there is the risk of fire.
At the point where the unprotected cable contacts the airframe metal there will undoubtedly be sparks, which will ignite anything
in the vicinity and the cable itself will get so hot that more insulation will be melted from the cable,
making the situation worse.
Although there are a number of protection devices used in aircraft electrical systems, we'll look at only two of them.
These are fuses and circuit breakers.
The fundamental difference between fuses and circuit breakers is the time it takes for each of them to operate from
the moment of maximum fault current flowing in the circuit.
A fuse normally breaks a circuit before full fault current is reached, whereas a circuit breaker operates to break the
circuit after full fault current is reached.
The circuit breaker can also function in certain circumstances as a switch, opening and closing a circuit as required.
The most common type of fuse in use on light aircraft is the cartridge fuse. It consists of a tubular
glass or ceramic body with the fuse element running through the centre, connected to two brass end caps.
The fuse operates, or blows, when the current flowing through is sufficient to melt the fuse element.
The time this takes varies inversely with the current.
All fuses are rated at a specific current value, that current being the current the element will carry continuously without
unduly heating up or deteriorating.
The rating of a fuse in a particular circuit is such that it is not less than the normal current
flowing in that circuit, but will blow at a current level below the safety limit of the equipment or cable
used.
For this reason, only the specified fuse rating should be used in a particular circuit.
A fuse that is blown may be replaced with another of the correct rating, but only once.
If the replacement fuse blows when the circuit is reactivated, then there is obviously a defect in the system and
the fuse must not be changed again until the circuit has been investigated.
Fuses which have a greater rating than that specified for a particular circuit must never be used.
The consequences will most probably include greater damage being caused and may be an electrical fire.
Circuit breakers are fitted, as are fuses, to protect equipment from overload or fault conditions.
They incorporate a heat sensitive tripping device and a manually operated trip reset switch.
The type of circuit breaker shown in the diagram can be rated from between 5 to 45 amperes.
The push pull button is shown in the operated position.
The white marker band is specifically there to draw attention to the fact that the circuit breaker has blown.
A circuit breaker which has operated, or popped, may be reset just once.
Do not be tempted to reset it a second time.
If it has popped twice then you may be sure that an electrical fault is present, and resetting the circuit
breaker twice will only exacerbate it.
This diagram shows a simplified alternator control circuit.
For the purpose of this lesson we will assume that our aircraft is fitted with an alternator.
From the point of view of the pilot, generator and alternator control circuits are very similar.
The only difference worth noting is the rectifier shown here.
This changes the alternating current output of the alternator to direct current, a procedure which is unnecessary with the generator
because its output is direct current.
At the top of the diagram you can see the buzz bar, which we previously said was a distribution point
for the output of the alternator or generator.
To the left is the field circuit, and to the right is the power circuit.
The field circuit carries a relatively small current. The current flows through the field windings and is used to generate
a magnetic field inside which the armature of the alternator rotates.
The field circuit of this alternator is protected by a 5A circuit breaker.
Just below that on the diagram is the alternator master switch, which you can see in the photograph is on
the left of the row of switches.
Immediately adjacent to the alternator master switch is the battery master switch.
These two switches are usually mechanically interlocked so that the alternator cannot be operated without the battery.
If either the field circuit breaker or the alternator master switch is opened then the alternator field becomes inoperative and
the alternator will not produce any output.
Also in the field circuit is the over volts relay. A relay is just an electrically operated switch.
This relay will operate to open the field circuit if the alternator output voltage rises above 16.5 volts.
The effect will be the same as if the field circuit breaker or alternator master switch had been opened. The
alternator will not produce any output.
In this particular aircraft type the over voltage relay can be reset by switching off the alternator master switch for
about 2 seconds and then switching it back on again.
Adjacent to the alternator is the voltage regulator. As its name implies it controls the voltage output of the alternator,
in this case at between 14.5 to 16 volts.
The power circuit passes the whole of the output of the alternator to the buzz bar, where it is then
distributed to the individual components in the aircraft.
The rectifier is necessary, as we said earlier, to convert the alternating current output of the alternator to direct current.
The load meter measures the total current flowing from the alternator to the buzz bar.
This illustration shows a simplified electrical wiring diagram of a typical light aircraft.
Notice that in this aircraft all of the loads but two are protected by circuit breakers of varying amperage.
The exceptions are the power side of the starting circuit and the battery.
Although the power side of the starter motor circuit is not protected, the control part of that circuit is.
Note that the starter switch gets its supply from the buzz bar via a 15 amp circuit breaker.
A starter motor will take considerably more than 15 amps when it is first selected upon engine start.
In fact the initial load on the battery will exceed this figure by more than a factor of 4.
That sort of current flow would cause a circuit breaker to pop open in an instant, and if it was
flowing through the starter switch it would either melt its contacts or weld them together.
Neither prospect being particularly appealing.
To prevent either of these occurrences, the starting current is taken from the battery and passed to a starter relay.
The starter relay is remotely controlled by the starter switch.
When the starter switch is selected, it sends a supply to energise the coil of the starter relay.
The starter relay has large contacts which are capable of easily passing the very high current required by the starter
motor, so the starter motor rotates the engine.
Try pressing the starter switch now, and note how the supply is passed through the starter relay.
Once the engine has started, the current that has been drained from the battery must be replaced by the alternator.
We've already stated that initially this current flow, which is shown on the load meter, will be quite high.
But we also stated that it should quickly reduce as the battery becomes recharged.
If no loads at all are selected, then the only current flowing through the load meter will be the battery
charging current.
Even when the battery is fully recharged, residual charging current will still be about 2A.
For night flying, for instance, add this 2A charging current to the other loads placed on the alternator and a
reading of approximately 32A would be representative.
If, on the other hand, the load meter reading ever drops to zero in flight, then this probably indicates that
the alternator has failed.
This occurrence will be covered in much greater detail later on in this lesson.
So far we've only shown the load meter, which has the zero position on the left side of the dial.
Another type of ammeter has the zero in the centre of the dial, as shown here, and is referred to
as a centre zero ammeter.
Notice that the centre zero ammeter is differently situated in the circuit from the load meter.
Whereas the load meter was capable of reading the whole of the output of the alternator, the centre zero ammeter
reads the current flowing into and out of the battery.
Prior to engine start, with perhaps the navigation lights, the radio, instrument panel lights and the electrical fuel pump selected
on, the centre zero ammeter will show by the fact that the needle is indicating in the negative portion of
the dial that the battery is discharging.
Just after start, when the battery is some way to having recovered its charge, the centre zero ammeter needle will
indicate in the positive portion of the dial.
This tells us that the alternator is capable of supplying all of the loads and the battery's charging current.
If, on the other hand, with the engine running, the centre zero ammeter needle is well into the negative portion
of the dial, it shows us that the alternator is incapable of supplying demand and that the battery is discharging.
In this situation it would be wise to start shedding loads, in other words, start switching off those services which
are unnecessary, until the ammeter needle once again indicates in the positive portion of the dial.
This will then indicate that there is a flow of charge current into the battery.
Most malfunctions of the aircraft electrical system will be indicated to the pilot by either the illumination of warning lights
or by the readings on the load meter or centre zero ammeter.
If the aircraft has an enunciator system, then it's likely that this includes an alternator failure warning light.
A test button will be incorporated into the enunciator system to check the light bulb filaments.
The ammeter readings have already been covered, but it may help to cover them once more from the point of
view of the indications of malfunctions.
First we'll talk about the type of ammeter which has the zero on the left side of the dial, the
load meter.
This will indicate alternator failure by the needle going to zero.
If the load meter reading appears high after the time the battery would normally have recharged itself subsequent to engine
start, this may indicate that the battery has an excessive charging rate.
An excessive charge rate will cause the battery to start losing some of the active material off its plates.
This will cause serious and lasting damage to the battery.
The high charge will also cause the battery to get very hot, possibly so much so that the electrolyte will
evaporate, exposing the plates to the air which once again will cause them damage.
Bear in mind that the high charge rate could perhaps be brought about by a faulty voltage regulator.
If this is the case then all of the aircraft equipment will be at risk of becoming overheated and impaired,
especially heat sensitive components like the radio and navigation equipment.
The centre zero ammeter will indicate alternator failure by the needle showing a constant heavy discharge.
If the needle stays just in the negative portion of the dial for any length of time then this is
characteristic of the situation where the alternator is incapable of supplying all the loads of the battery at the same
time.
This will require that some of the loads are switched off, in other words some judicious load shedding must be
done, otherwise the battery will eventually be completely discharged.
The following drills are not to be considered as being representative of emergency drills for any particular aircraft type.
They are purely general recommendations of the type of action which should be considered if an alternator should fail.
In the event of either the load meter showing zero or alternatively the centre zero ammeter indicating a heavy discharge
and or the alternator failure warning light being on, the following actions should be considered.
Initially load shed, that is judiciously select off any electrical, radio and navigation services not vital to the safe operation
of the aircraft.
If the alternator failure warning light is not illuminated then use the press to test facility to check its filament.
Next check the field circuit breaker to see if it has tripped. If it looks normal consider tripping and resetting
it anyway.
If your aircraft system has over voltage protection and the circuit breaker has not tripped these indications may be evidence
that the over voltage relay has tripped.
Switch off the alternator master switch and leave it off for about 2 seconds. Try that now.
Switch the master switch back on again.
Now check the load meter or alternatively the centre zero ammeter and also the alternator failure warning light.
If the indications are that the alternator output has been restored then restore the services singly in order of their
importance to the safe operation of the aircraft.
Do this relatively slowly and deliberately. Allow sufficient time between each service selection for the load to be taken up
by the alternator.
Should the alternator fail again then the drill should be repeated but this time do not select the faulty service.
If your attempt at restoring the alternator is unsuccessful then it would be advisable to reduce the electrical load as
much as possible and land as soon as possible.
Remember any radio transmissions you make will seriously reduce the amount of charge left in the battery.
Most modern light aircraft use engine driven suction pumps to generate a vacuum which induces an airflow into the gyro
operated instruments.
The airflow is directed over buckets formed in the rim of the rotor of the instrument gyros and this spins
them round at very high speed.
The instruments which are usually powered by this system are the attitude indicator and the heading indicator, sometimes called the
directional gyro.
The vacuum system consists of a means of generating a vacuum, a method of controlling the vacuum, a filter to
clean the air being sucked through the instruments,
and the necessary pipework to join all the components together.
Methods of generating the vacuum range from connecting the system to the inlet manifold of a normally aspirated piston engine,
to a venturi placed in the airstream, through to either engine driven or electrically driven vacuum pumps.
Older aircraft may have a venturi tube to generate the vacuum.
The venturi tube works on the principle that when it's placed in the airstream its effect is to accelerate the
air as it passes through.
When the air accelerates its pressure drops.
At the narrowest point of the venturi the speed of the air is greatest and its pressure lowest.
Once through the narrowest point the air decelerates and its pressure rises once more.
Placing the suction tube at the point of lowest pressure makes use of the partial vacuum thus generated.
It must be appreciated that this system will only work reliably when sufficient air is flowing through the venturi.
This may not happen for some time after the aircraft has become airborne.
The elapsed time will allow the partial vacuum to become effective in driving the gyros fast enough to give the
correct indications on the instruments.
A few aircraft are fitted with an inlet manifold suction system.
This works on the principle that whenever a normally aspirated piston engine is running there is a depression created in
the inlet manifold.
This depression is greatest when the engine is just ticking over with the throttle almost fully closed and gets less
as the throttle is opened.
The partial vacuum thus created at low to medium engine speeds should certainly be sufficient to drive the instrument gyros
at high speed.
Some early light aircraft, those built in the 1960s, were fitted with wet vacuum pumps.
They used metal vanes which were lubricated with engine oil.
Wet vacuum pumps wore out slowly and gradually and were very reliable.
Most modern light aircraft since about 1970 have used what are termed dry vacuum pumps to power their suction system.
They are called dry because they use self-lubricating graphite vanes which rotate within an eccentric aluminium cavity.
Dry vacuum pumps are generally very reliable however when they fail they do so catastrophically and without warning.
To ensure that a failure cannot damage the engine's accessory drive the pump drive incorporates a shear coupling
which will break if the pump's input torque exceeds its normal operating figure by any significant amount.
If the drive does shear then of course the gyro driven instruments will cease to function.
To cover for this eventuality some aircraft have an electrically driven auxiliary vacuum pump
which will provide an independent backup in the event of an engine driven pump failure.
Air driven gyro instruments are designed to operate with a pressure differential of about 5 inches of mercury.
Modern engine driven vacuum pumps are designed to produce a large surplus of airflow through the instruments
even when the engine is only running slowly on the ground.
At engine cruise RPM the pumps increase their capacity by a factor of 8 or more
and as a consequence if the suction was uncontrolled the gyros would be driven far too fast.
A constant airflow through the gyros is maintained by the pressure regulator
which admits sufficient air to leak into the system downstream of the gyros
to limit the pressure differential across them to about 5 inches of mercury.
The pressure regulator in most systems is adjustable.
The vacuum or suction gauge is connected to read the pressure differential across the gyro driven instruments.
Most suction gyro instruments require that the gauge reads 5 inches of mercury to ensure that they operate correctly.
Ambient air enters the system through a vacuum filter.
This ensures that the gyros are fed with only clean air that has had dirt and other contaminants removed from
it.
When the filter is first fitted to the system there is almost no pressure drop across it.
Dry vacuum pumps are very vulnerable to contamination by liquids.
The graphite veins are designed to operate in an absolutely dry environment
and if any form of liquid, for instance water, oil or engine cleaning fluid gets into the pump then the
veins will be destroyed.
Another form of contamination which will quickly destroy a dry pump is carbon.
When a pump fails it will do so suddenly, generating a cloud of carbon fragments which can lodge in the
outlet pipes of the pump.
If the system is not meticulously cleaned when a replacement pump is being fitted then there is a danger that
these fragments will fall back into it.
If this is the case then the replacement pump will not last long.
Pump failure can also be brought on through it becoming hot because it's having to work too hard to produce
the required suction level.
This can be the result of a partially blocked filter or perhaps a hose in the system collapsing or having
a kink in it.
If the pressure regulator is adjusted to bring the system pressure back to its normal operating level then the pump
will be working harder to produce the same results in the cockpit.
A pump that is working too hard will run hot and will fail prematurely.
If there is a system malfunction there are two indications in most aircraft cockpits to warn the pilot of vacuum
system failure.
They are the suction gauge and the low vacuum warning light.
A further clue may be obtained from the operation or otherwise of the gyro driven instruments.
If the suction gauge is showing system failure by indicating either low or no suction and at the same time
both the gyro compass and the attitude indicator appear to be functioning then the problem would seem to lie with
the suction gauge itself.
Is the gyro compass telling you that you are not maintaining a steady heading and is the attitude indicator showing
wings level?
If the answer is yes to either or both of these questions then you have to determine where your problem
lies.
If both instruments appear to be giving erroneous indications then it is a fair assumption that the suction system has
failed.
This of course should already have been flagged up to you by the suction gauge and the low vacuum warning
light.
If on the other hand only one or the other of the attitude indicator or heading indicator appears to be
telling lies then your flying instructor should have given you the ammunition to enable you to determine the correct course
of action.
If you determine that your engine driven vacuum pump has failed and your aircraft is fitted with an auxiliary vacuum
pump now is the time to select it on.
A couple of words of caution however. First be aware that the electrical load that these pumps place on your
system is quite substantial.
It may be that you will have to consider load shedding of electrical services in order that you do not
compound your misery with an alternator failure as well.
Secondly the strength of the magnetic field generated by the electric motor in the pump is such that it can
cause compass errors of plus or minus 10 degrees.
It would be a good idea to note the magnetic compass reading before and after selection of the auxiliary vacuum
pump.
Your aircraft may have inlet manifold suction fitted as a reserve for the normal vacuum system.
It should be selected as a backup as soon as you are sure of the indications that your engine driven
vacuum pump has failed.
Remember though that at high engine power settings the depression in the inlet manifold may not be sufficient to create
the required pressure differential to drive the aircraft instruments.
It may be necessary to reduce the throttle setting until your suction gauge indicates that you have regained enough vacuum.
In the unfortunate event that you do suffer a failure of the suction system and you have no backup system
then you'll have to revert to the limited panel flying technique taught to you by your flying instructor.
The Pitot static system supplies both pitot and static pressure for the airspeed indicator, the altimeter and the vertical speed
indicator.
Pitot and static pressure are usually sensed by a pitot head situated where it's not likely to be affected by
turbulent air flow caused by projections from the aircraft's skin.
Some aircraft have static vents which are separate from the pitot head. These are usually fitted in the side of
the fuselage.
Pitot and static pipelines within the wing and fuselage carry the pressure to the gauges on the instrument panel.
An alternate static source is usually provided, the control valve for which will be mounted somewhere close to the instrument
panel.
With an unpressurised aircraft, if the alternate static source is from within the cockpit then the pressure sensed will normally
be lower than normal outside static pressure due to aerodynamic suction.
It's inevitable that water will enter the pitot static system and it is just as inevitable that should that water
enter the barometric instruments associated with the system then they will suffer damage.
To prevent the water getting too far into the system, drain valves are fitted at the lowest point in the
pipeline.
Part of the pre-flight check should involve opening these valves to allow any moisture to escape.
Ensure that the valves are closed properly after the check.
An aircraft at rest on the ground in still air is subject to normal atmospheric pressure which bears equally on
all parts of the aircraft.
This ambient pressure is known as static pressure.
An aircraft in flight, while still subject to the static pressure at its flight level, experiences an additional pressure on
the leading edges due to the resistance of the air to the aircraft's movement.
This additional pressure is called dynamic pressure and its value depends both on the speed of the aircraft through the
air and on the density of the air.
The leading edges therefore encounter a total pressure consisting of static plus dynamic pressures.
This total pressure is also known as pitot pressure.
An open-ended tube, parallel to the longitudinal axis of the aircraft, is used to sense the total pressure, that
is static pressure plus dynamic pressure.
This device is a pitot tube, which is mounted in a pitot head.
The open end of the tube faces into the moving airstream, while the other end leads to the airspeed capsules
in the airspeed indicator.
The moving airstream is thus brought to rest in the tube, so generating the extra dynamic pressure, which together with
the static pressure already in the tube, provides the required total or pitot pressure.
The static and pitot sources may be combined in one pressure head, the static tube surrounding the pitot tube, with
separate pressure lines leading to the pressure instruments.
An electric anti-icing heater coil is usually incorporated, which is controlled from a switch in the cockpit.
It will be appreciated that, if due to turbulent airflow in the region of the pitot static heads, the pressures
sensed are not truly representative of the pitot and static pressures, the pressure-dependent instruments will not read correctly.
The error involved is called position error or, alternatively, pressure error.
Approximately 95% of the position error associated with a combined pressure head is produced by turbulence about the static
head.
Because of this, the static vent was introduced as a source of static pressure instead of the static head.
Pitot pressure then being sensed by a simple pitot head.
About 90% of the combined pressure head position error is eliminated by use of a separate pitot head and
static vent.
A separate static vent is shown here.
There is usually some place on the airframe, usually on the side of the fuselage, where true or nearly true
static pressure obtains over the whole speed range of the aircraft.
Earlier in this lesson, we learnt that whereas an aircraft on the ground in still air is subject only to
static pressure, the leading edges of the aircraft in forward flight are subject to an additional pressure which we called
dynamic pressure.
This results in a total pressure on the leading edges of dynamic pressure plus static pressure, which we called pitot
pressure.
Expressed as a formula, it looks like this.
Pitot pressure equals dynamic pressure plus static pressure or P equals dy plus s.
The pitot head senses pitot pressure and the static vent senses static pressure.
These two pressures are fed to the airspeed indicator, which is a differential pressure gauge, and it measures the dynamic
pressure which is the difference between pitot and static.
Dynamic pressure is a measure of airspeed because dynamic pressure equals a half rho V squared, where V is true
airspeed and rho is the density of the surrounding air.
Thus the airspeed indicator measures airspeed by measuring dynamic pressure, displaying the result, usually in knots, on a suitably calibrated
scale.
In principle, the simple airspeed indicator can be considered as an airtight box, divided by a flexible diaphragm, with pitot
pressure fed to one side of the diaphragm and static pressure to the other side, as shown here.
The pressure difference across the diaphragm is dynamic plus static minus static, which is dy, the dynamic pressure.
Accordingly, the diaphragm is deflected by an amount proportional solely to the dynamic pressure, its movement being transmitted by a
system of levers to the indicating needle on the face of the airspeed indicator.
Note that static pressure is common to both sides of the diaphragm, and so does not influence diaphragm movement.
In practice, as shown here, the static pressure is fed into the hermetically sealed instrument case, the pitot pressure being
piped to a thin metal capsule capable of expansion and contraction.
Note that the pressure differential between the inside and outside of the capsule is exactly the same as it was
with the diaphragm in the previous diagram.
Expansion or contraction of the capsule will therefore be proportional to the changes in dynamic pressure produced by changes of
airspeed.
The capsule movements are transmitted via temperature compensated magnifying linkage to the pointer indicating airspeed on the face of the
airspeed indicator.
DYNAMIC PRESSURE
Previously we used the formula dynamic pressure equals a half rho V squared.
From this we can see that dynamic pressure depends not only on the speed of the aircraft, but also on
the air density, rho.
This density varies with temperature and pressure and thus with altitude.
The airspeed indicator is calibrated to read true airspeed for the air density of 1225 grams per cubic metre,
which would be produced by the international standard atmosphere mean sea level pressure of 1013.25 millibars and a temperature
of plus 15 degrees Celsius in dry air conditions.
No allowance is made in the calibration for the change in density which occurs with change of altitude.
It follows that, even if there were no other errors, the airspeed indicator could only indicate the true airspeed when
international standard atmosphere mean sea level air density existed at its flight level.
That's to say when the aircraft was flying in air having a density of 1225 grams per cubic metre.
This density value can only be found at or close to sea level, depending on how much the actual conditions
deviate from standard.
Manufacturing imperfections and usage result in small errors which are determined on the ground under laboratory conditions by reference to
a DYNAM instrument.
A correction card can be produced for the speed range of the instrument.
We heard earlier that position error is alternatively known as pressure error.
Position error arises mainly from the sensing of incorrect static pressure.
Position errors throughout the speed range are determined by the aircraft manufacturer during the test flying programme for a particular
aircraft type.
It's not unusual to compile a joint correction card for position and instrument errors and place it in the aircraft
near the ASI content.
To summarise then, the indicated airspeed plus or minus instrument error correction and position error correction equals the calibrated airspeed.
Unless the air around the aircraft is at the calibration density of 1225 grams per cubic metre, which can only
occur near sea level, the airspeed indicator cannot correctly indicate true airspeed.
From our formula, we know that dynamic pressure is proportional to density.
So at altitude, where density is less, the dynamic pressure generated by a given true airspeed will be less than
for the same true airspeed in flight at sea level.
Airspeed indicator capsule expansion will be proportionately less and the speed indicated will be less than the true speed.
The difference between indicated airspeed and true airspeed is called density error.
This density error will cause the airspeed indicator to under-read at altitude, whereas if you get to fly below
mean sea level, it will over-read.
Use the mouse to move the aircraft higher or lower and note the change in true airspeed.
It's possible on some navigation computers, like the CRP5, to be able to correct for density error, computing true airspeed
from the calibrated airspeed which we mentioned earlier.
This computation allows for the height of the aircraft above the calibration 1013.25 mbar pressure level and also for
any temperature deviation from standard conditions, summarising the true airspeed equals the calibrated airspeed plus or minus the correction for
density error.
Some airspeed indicators incorporate coloured markings on the dial. These range markings consist of coloured arcs and radial lines.
The white arc denotes the flap operating range, from stall at maximum all-up weight in the landing configuration, full
flap, landing gear down, wings level, power off, up to VFE, maximum flaps extended speed.
The green arc denotes the normal operating speed range, from stall speed at maximum all-up weight, flaps up, wings
level, up to VNO, normal operating limit speed or maximum structural cruising speed, which should not be exceeded except in
smooth air.
Operations at indicated airspeeds in the green arc should be safe in all conditions including turbulence.
The yellow arc denotes the caution range, which extends from VNO, normal operating limit speed, up to VNE, the never
exceed speed.
The aircraft should be operated at indicated airspeeds in the caution range only in smooth air.
A red radial line denotes VNE, the never exceed speed.
If the pitot head becomes blocked the airspeed indicator reading will, in general, remain unchanged.
In level cruise a blockage will lock in the previous pitot pressure and any change in actual airspeed will not
be registered.
If altitude is changed with a blocked pitot head and a clear static source, the indicated airspeed will increase during
a climb because the pressure locked inside the capsule remains constant while the static pressure of the air surrounding the
capsule decreases.
Conversely, the indicated airspeed decreases during a descent with a blocked pitot head.
A leak in the pitot pipeline will cause the airspeed indicator to under-read.
The effect of leaks in the static pipeline is normally to make the airspeed indicator over-read.
This is because almost inevitably the pressure surrounding the pipeline is lower than static because of aerodynamic suction.
A static head is more exposed to icing conditions and is therefore more likely to become obstructed than is a
static vent.
A blocked static source during descent will mean that the old, higher altitude static pressure surrounding the capsule will be
lower than it should be so that if the pitot pressure is normal the airspeed indicator will over-read.
This could be dangerous in that the aircraft is nearer the stall than the airspeed indicator is showing.
Note that a climb with blocked static source and normal pitot air will result in the airspeed indicator under-reading.
The following checks of the airspeed indicator and pressure supply system should be made before flight.
If they are fitted to your aircraft, pressure head covers and static vent plugs must be removed and stowed aboard
the aircraft.
The pitot tubes, holes or slots in the static heads and or static vents should be checked to be free
from obvious obstructions such as insects.
Next, if one is fitted, the pitot head heater must be checked to see that it is operative.
The instrument dial glass must be clean and undamaged.
Finally, the instrument should indicate airspeed in the correct sense shortly after starting the take-off run.
The pressure altimeter is a simple, reliable pressure gauge calibrated to indicate height.
The pressure at a point depends on the weight of the column of air which extends vertically upwards from that
point to the outer limit of the atmosphere.
The higher an aircraft is flying, the shorter is the column of air above it, and consequently the lower is
the atmospheric pressure at the aircraft.
In other words, the greater the height, the lower the pressure, and by measuring the pressure, the altimeter measures height.
Unfortunately, the relationship between pressure and height is not a linear one, so that calibration of the altimeter scale is
not a simple matter.
The situation is further complicated by high and low pressure weather systems, which produce pressure differences in the horizontal plane.
Furthermore, the temperature of the air at the surface and the temperature lapse rate in the air above vary considerably,
and this affects pressure.
With all the variables and changing relationships that abound in the atmosphere, it becomes necessary to assume certain average or
standard conditions.
We can then base the altimeter calibration formulae on these, and then apply corrections appropriate to the deviations from standard
conditions which occur with position and time.
The conditions used for calibration are usually those assumed for the International Standard Atmosphere, or the ISA.
Several assumptions are made in constructing the International Standard Atmosphere.
The first is that at mean sea level the pressure is 1013.25 mbar, the temperature is plus 15 degrees
Celsius, while the density is 1225 g per cubic meter.
The second is that from mean sea level up to 11 km, or 36,090 ft, the temperature falls at
6.5 degrees Celsius per kilometer, or 1.98 degrees Celsius per 1,000 ft.
The third assumption is that from 11 km to 20 km, or from 36,090 ft to 65,617 ft,
the temperature is constant at minus 56.5 degrees Celsius.
And finally, from 20 km to 32 km, or from 65,617 ft to 104,987 ft, the temperature rises
at 1 degree Celsius per kilometer, or at 0.3 degrees Celsius per 1,000 ft.
Having made these assumptions, the pressure corresponding to any given level in the ISA can be calculated from the calibration
formula.
Graphs or tables can be produced showing height in terms of pressure under standard conditions.
These tables can be used for the manufacturer's calibration of the altimeter scale.
Basically, the laboratory calibration consists of applying a series of pressures to the altimeter and checking that the instrument indicates
the respective levels, which correspond to these pressures in the ISA.
Any discrepancies, if within certain agreed tolerances, would be listed over the operating height ranges as instrument errors.
Static pressure is fed into the case of the instrument from the static source. As height increases, static pressure decreases
and the capsule expands under the control of a leaf spring.
Conversely, as height decreases, static pressure increases and the capsule contracts.
A mechanical linkage magnifies the capsule expansion and converts it to a rotational movement of a single pointer over the
height scale.
The linkage incorporates a temperature compensating device to minimise errors caused by expansion and contraction of the linkage and changes
in spring tension due to fluctuations in the temperature of the mechanism.
This animation shows how the linkage works, but the actual arrangements are much more complex.
The simple altimeter can have a setting knob which is geared to the pointer.
With this knob, the pointer may be set to reach zero with the aircraft on the ground, so that when
airborne, the altimeter indicates the approximate height above aerodrome elevation.
Alternatively, the pointer can be set, before flight, to the aerodrome elevation, so that when airborne the instrument shows approximate
height above mean sea level.
Most aircraft are now equipped with a sensitive type of altimeter similar to the one shown here.
The principle of operation is similar to that of the simple altimeter, but there are some refinements.
For instance, a bank of two or three capsules gives the increased movement necessary to drive any counters and pointers.
Also, jewelled bearings are fitted in the linkage, reducing friction, and thus the associated lag in indication.
Finally, and probably most importantly, a variable datum mechanism is built in.
This, with the aid of a setting knob, enables the instrument to be set to indicate height above any desired
pressure datum.
The variable datum mechanism can be used to set the desired pressure level, say 1005 millibars, on a pressure subscale
on the face of the instrument.
As the pilot turns the setting knob, the altimeter reading changes until, when the procedure is completed with the subscale
showing the desired 1005, it indicates the aircraft's height above this pressure level.
A change in altitude or surface pressure has no direct effect on the reading of the subscale. It only changes
when the pilot turns the setting knob.
British altimeters have a subscale setting range between 800 to 1050 millibars.
The three pointer instruments shown here give a much more sensitive indication of height and change of height, but they
do suffer from the severe disadvantage that they can be easily misread.
It's not difficult for the pilot to make a reading error of 10,000 feet, particularly during a rapid descent
under difficult conditions with a high workload.
Accidents have occurred as a result of such misreading.
Various modifications to the pointers and warning systems have been tried with the object of preventing errors, including a striped
warning sector which appears as the aircraft descends through the 16,000 foot level.
However, the greatest advance has been the introduction of the counter pointer altimeter, shown here on the right.
This gives a much more positive indication than either of the three pointer dials. Compare the simplicity of the indication.
The transition altitude is the altitude at or below which the vertical position of an aircraft is expressed and controlled
in terms of altitude.
In the United Kingdom its value is commonly 3,000 or 4,000 feet.
The transition level is the lowest flight level available for use above the transition altitude. At and above transition level
vertical position is expressed as a flight level.
The transition layer is the airspace between transition altitude and transition level.
When climbing through it, the aircraft's vertical position is expressed in terms of flight level, when descending through it in
terms of altitude.
Flight levels are surfaces of constant pressure related to the standard pressure datum of 1013.25 millibars.
Flight levels are separated by specified pressure intervals.
In the United Kingdom these intervals are 500 feet between transition level and flight level 245, while from flight level
250 upwards they are 1,000 feet.
A flight level is expressed as the number of hundreds of feet which would be indicated at the level concerned
by an ISA calibrated altimeter set to 1013.25 millibars or 29.92 inches of mercury.
For example, with 4,500 feet indicated, the flight level would be flight level 45.
It should be remembered that the altimeter indicates height above the pressure level set on the subscale.
The three settings normally used on the subscale are shown on these buttons. Select them to find the definitions of
each.
QFE is the aerodrome level pressure, which when set on the subscale will cause the altimeter of an aircraft on
the ground to read zero, assuming there is no instrument error.
In flight, with QFE set, the altimeter will indicate height above the aerodrome QFE reference datum, provided ISA conditions obtain
between aerodrome level and the aircraft and there are no other altimeter errors.
In practice, QFE is used mainly for circuit flying and gives a good indication of height above the aerodrome.
QNH is used mainly when flying below transition altitude. It's an equivalent mean sea level pressure, calculated by air traffic
control from the aerodrome level pressure.
With QNH set on the subscale, the altimeter of an aircraft on the aerodrome indicates aerodrome elevation, that is, the
height above mean sea level.
In flight, the altimeter will indicate altitude, but this will only be the true altitude, the height of an aircraft
as measured from mean sea level, if the mean temperature in the column of air beneath the aircraft is the
same as in ISA conditions.
When 1013.25 mbar is set on the subscale, the altimeter shows the pressure altitude or pressure height. This setting
is used in the United Kingdom above the transition altitude or transition level.
Select any of these buttons to learn more about the main altimeter errors.
With many types of altimeter, the response to change of height is not instantaneous. This causes the altimeter to under
-read in a climb and over-read in a descent. The lag is most noticeable when the change in altitude
is rapid and prolonged.
Manufacturing imperfections, including friction in the linkage, cause errors throughout the operating range. The errors are kept as small as
possible by adjustments within the instrument, and the calibration procedure ensures that they are within permitted tolerances.
Residual errors may be listed on a correction card.
Position error is largely due to the inability to sense the true static pressure outside the aircraft as we mentioned
earlier. The error is usually small.
Manoeuvre induced error is caused mainly by transient fluctuations of pressure at the static vent during change of pitch attitude.
Even with no other errors at all, the pressure altimeter will not indicate the true altitude, the height above mean
sea level, unless the surface temperature and lapse rate of the column of air are those which were assumed in
the calibration of the altimeter.
Where the temperature at your cruising level deviates from the international standard atmosphere, an approximate correction can be made with
most navigational computers. The correction can only be approximate since temperatures in the rest of the column of air are
not known.
When flying in air which is colder than the international standard atmosphere, which will have an air density greater than
international standard atmosphere at that altitude, the altimeter will over-read.
When flying in air which is warmer than the international standard atmosphere, which will have an air density lower than
international standard atmosphere at that altitude, the altimeter will under-read.
If the static source becomes blocked, the altimeter will not register any change in height. The height at which the
blockage occurred will still be indicated, regardless of any climb or descent.
The following checks of the altimeter should be made before flight. The instrument dial glass must be clean and undamaged
and the correct altimeter setting must be set on the subscale.
Air traffic will pass you the correct altimeter setting and once you have it set you can check that the
altimeter reads within plus or minus 50 feet of the apron elevation.
The apron elevation should be displayed in the airfield operations centre or, alternatively, it may be marked on the aeronautical
chart you are using.
When an aircraft departs from level flight, the static pressure changes. The vertical speed indicator measures the pressure difference between
each side of a metering unit.
In level flight, the pressures on each side of the metering unit are the same. During a climb or descent,
air fed to the capsule immediately responds to the change of atmospheric pressure, but the metering unit transmits this change
at a lower rate.
Static pressure is fed to a capsule enclosed within an airtight case. The case itself is also fed with static
pressure, but this feed comes through a restricted metering unit.
Thus, if the static pressure is changed, the pressure surrounding the capsule changes at a slower rate than that within
the capsule.
If the aircraft is climbing, the pressure in the capsule will be less than that in the case. The consequent
compression of the capsule is converted by a suitable linkage to a pointer indication of the rate of climb.
If the aircraft is descending, the pressure in the case will be less than the pressure in the capsule, causing
the capsule to expand, which will move the linkage to indicate the rate of descent.
Two types of presentation are available, a linear scale and a logarithmic scale. It will be obvious from comparing these
two diagrams that the presentation of a logarithmic scale is easier to read at the lower rates of climb and
descent.
Select any of the four buttons to learn more about the main errors that occur within the vertical speed indicator
and its associated equipment.
Instrument error is due to manufacturing imperfections.
If the static pressure is subject to position error, the vertical speed indicator will wrongly indicate a climb or descent
when speed is suddenly changed. This is most noticeable during take-off acceleration.
Any short-term fluctuations in pressure at the static vent during attitude changes will cause the instrument to indicate a
false rate of climb or descent.
Additionally, most vertical speed indicators have a small counterbalance weight included in the linkage, the inertia of which causes delays
in the indications of changes in vertical speed during manoeuvres.
The vertical speed indicator pointer takes a few seconds to steady because of the time taken to build up a
steady pressure difference on climb or descent.
There will also be a time lag on levelling out because of the time taken for the pressures to equalise.
This error is most noticeable after a prolonged climb or descent, especially at high rate.
To overcome the problem of lag, the instantaneous vertical speed indicator incorporates an accelerometer unit which responds quickly to a
change of altitude.
Select either descent or climb to see what happens in the instrument.
The piston in the vertical acceleration pump immediately falls in the cylinder and decreases the pressure in the capsule.
The capsule contracts and the pointer will give an instant indication of climb.
If the vertical acceleration ceases, after a few seconds the piston will slowly rise to its original position, but by
this time the correct differential pressure between the capsule and the case will have been set up and the correct
rate of climb will continue to be shown.
The piston in the vertical acceleration pump immediately rises in the cylinder and increases the pressure in the capsule. The
capsule expands and the pointer will give an instant indication of descent.
If the vertical acceleration ceases, after a few seconds the piston will slowly descend to its original position, but by
this time the correct differential pressure between the capsule and the case will have been set up and the correct
rate of descent will continue to be shown.
Any blockages of the static line or vent will cause the needle to return to zero.
If the supply of air to the vertical speed indicator is blocked, it's probable that the other pressure instruments, the
airspeed indicator and the altimeter, will also be affected.
The following checks of the vertical speed indicator should be made before flight.
The instrument dial glass must be clean and undamaged and the instrument should read zero.
In flight, the accuracy of the instrument may be checked against the altimeter and a stopwatch during a steady climb
or descent, and the instrument should indicate zero climb or descent when in level flight.
Any rotating body exhibits gyroscopic phenomena.
The Earth is a gyro, spinning about the imaginary axis between the geographic poles.
A child's top is a gyro, and so is an aircraft's propeller when it's turning, especially at high RPM.
The rotor which forms an aircraft instrument gyro may be little more than an inch in diameter, spinning at perhaps
25,000 RPM.
Gimbal rings are the supports for the rotor of a gyroscopic instrument. Gimbal rings are sometimes just called gimbals.
This diagram shows a spinning rotor mounted in two gimbal rings, the outer one being supported by a fixed frame.
The rotor itself is a metal disc which rotates about this axis, usually called the rotor spin axis.
The rotor shaft, or spindle, is supported by bearings in a ring called the inner gimbal.
The inner gimbal is in turn supported by bearings mounted inside the outer gimbal, which can rotate on bearings in
the frame.
If an external force, or more correctly torque, is applied to change the direction of the rotor axis, the gyro
resists angular movement in the plane of the torque applied, and instead moves in a plane at right angles to
that of the torque.
The resulting movement being called precession.
The rule of precession states that the gyro will precess in a direction at 90 degrees to the applied force,
measured around the circumference of the rotor in the direction of spin.
The force applied appears to have moved 90 degrees in the direction of spin.
A spinning rotor maintains its axis pointing in a fixed direction in space, unless it's subjected to an external force.
This property is called rigidity in space, or gyroscopic inertia.
There are three factors which affect the rigidity in space of a gyro. They are the rotor mass, the effective
radius at which the mass acts, and the speed of rotation.
To increase the rigidity of air-driven gyros, the rotor is manufactured from brass, and to increase the effective radius
at which the mass acts, the brass is concentrated around the rim of the rotor.
Finally, the higher the rotational speed of the gyro, the greater its rigidity in space.
If a gyro has only one gimbal ring, with consequently only two planes of freedom, it can be adapted for
use as a rate gyro, which will measure a rate of angular movement.
If the gyro frame is rotated in the plane in which the gyro has no freedom, the rotor will precess,
unless restrained, until its plane of rotation coincides with the plane in which the frame is being turned.
If this precession is restricted by a spring, as shown in this diagram, the resultant tilt of the rotor will
be a measure of the rate of angular movement of the instrument.
This instrument incorporates two measuring devices, both indicating on the same instrument face.
One of these, the rate of turn indicator, commonly shortened to turn indicator, uses a rate gyro to measure rate
of turn about a vertical axis.
The other, the slip indicator, is a very simple pendulous device which is used mainly to show whether or not
a turn is balanced, whether the angular bank is correct for the true aspian rate of turn, and if not,
to indicate the extent of slip or skid.
The turn indicator employs a rate gyro, which, having only one gimbal, has freedom about only two axes.
There is of course complete freedom of rotation about one of these, the rotor spin axis, which in level flight,
lies a thwartships.
There is restricted freedom about the longitudinal gimbal axis.
There is no freedom about the aircraft's vertical axis, so any talk applied about this axis, as in a turn,
will cause the gyro to precess.
Suction and electrically driven types of turn and slip indicator are available.
With the former, an engine driven pump or venturi tube is used to apply suction to the instrument case.
Replacement air enters via a filter and is directed by a jet at buckets cut in the periphery of the
rotor.
The rotor rotational speed is low compared with that of the directional gyro indicator and the artificial horizon.
This is because the gyroscopic property of precession is used to measure rate of turn, and so a high gyroscopic
rigidity is undesirable.
A damping system fitted to the gimbal reduces oscillation.
This may be the piston in cylinder type or an electromagnetic device.
Stops limit the ability of the gimbal to tilt to an angle corresponding to a turn of about 20 degrees
per second.
If the suction to an air driven instrument is inadequate, which can happen as a result of either high altitude,
or with a choked filter, or even perhaps a leaking suction tube,
gyro rigidity will be lowered as the gyro is under-speeding.
Consequently, the secondary precession needed to equal the aircraft turn can be generated by a smaller secondary torque.
This reduced torque will be produced by a smaller angle of gimbal tilt, and this means that the instrument will
under-read the turn rate.
Alternatively, if the gyro were to over-speed, by the opposite process it will over-read the rate of turn
that is being achieved by the angle of bank applied.
It is desirable that turns should be properly balanced, with no side-slip or skid.
The slip indicator gives a direct indication of the state of balance of the turn.
Early types of slip indicator employed a simple metal pendulum suspended in the instrument case,
its oscillations being controlled by a piston-in-cylinder damping device.
The modern version, a sketch of which is shown here, is usually a ball-in-tube inclinometer.
This comprises a solid ball in a curved tube containing liquid which damps out the unwanted oscillations.
The heavy ball behaves like a pendulum, with the centre of curvature of the tube acting as the effective point
of suspension.
The turn coordinator is an interesting development of the turn and bank indicator.
The primary differences between the turn and bank indicator and the turn coordinator are in the setting of the precession
axis of the rate gyroscope, and in the method of display.
The gyroscope is spring-restrained and is mounted so that the axis is about 30 degrees with respect to the
aircraft longitudinal axis,
thus making the gyroscope sensitive to banking of the aircraft as well as to turning.
A turn is normally initiated by banking the aircraft.
The gyro will precess and this in turn will move the aircraft symbol to indicate the direction of bank and
enable the pilot to anticipate the resulting turn.
The pilot then controls the turn at the required rate by alignment of the aircraft with the graduations on the
instrument dial.
The rate of turn will depend on the instrument in use, either as a rate one turn, 3 degrees per
second, or any other rate dependent on instrument design.
To achieve a balanced rate of turn the ball still has to remain central.
The annotation No Pitch Information on the indicator scale is given to avoid any confusion regarding pitch information,
which might result from the similarity between this presentation and the presentation of the artificial horizon or attitude indicator.
The artificial horizon, or attitude indicator as it is sometimes called, provides the pilot with information in terms of the
aircraft's attitude both in pitch and roll.
It is a primary instrument, replacing the natural horizon in poor visibility.
The attitude display consists of a miniature aircraft shape or gull wing painted or engraved centrally on the inside of
the glass face of the instrument,
and therefore fixed to the instrument case and the actual aircraft.
Behind this representation of the aircraft is the horizon bar, linked to the gyro in such a way that the
bar is gyro-stabilised parallel to the true horizon.
The artificial horizon may be suction or electrically driven. It is also known as a gyro horizon and attitude indicator.
The artificial horizon uses an earth gyro in which the spin axis is maintained in or tied to the vertical
by earth's gravity.
This means that the plane of the rotor rotation is horizontal, thus it provides the stable lateral and longitudinal references
required.
An engine driven suction pump, or venturi tube in sunlight aircraft, is used to create a suction of about 4
or 5 inches of mercury in the instrument case of the air driven artificial horizon.
Replacement air, drawn in by this suction via a filter, is ducted through the outer and inner gimbals to enter
the rotor case as a jet, which spins the rotor at up to 15,000 rpm.
After driving the rotor, the air passes into the instrument case through slots at the base of the rotor housing.
The amount the case can move relative to the gyro is controlled by fixed stops. With older designs typical limits
are plus or minus 60 degrees in pitch and 110 degrees each way in roll.
In modern instruments there is complete freedom in roll and up to 85 degrees plus or minus in pitch.
If the limits are exceeded the gyro topples, giving violent and erratic movements of the horizon bar.
Unless a fast erection system is incorporated, accurate indications will not be obtained until the gyro has re-erected itself
over a period of 10 to 15 minutes.
The rotor assembly is made very slightly bottom heavy in order to keep down the time taken for initial erection
when the gyro is first started up.
But a complex control system is required to maintain the rotor axis vertical in flight.
Use your mouse to move the virtual control column, the black dot in the lower centre of the screen, to
see what effect its movement has on the artificial horizon and the altimeter.
Although this simulation is realistic in roll, there are limitations in the method of its production which prevent it showing
complete accuracy in pitch.
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Before flight, check that the horizon bar takes up a laterally level position with the correct pitch indication for the
aircraft type, and that this indication is maintained when taxiing.
If a caging device is fitted, the instrument should be uncaged at least 5 minutes before take-off to ensure
that the rotor axis has had time to reach alignment with the true vertical.
In flight, the artificial horizon should give an immediate and correct indication of any change in pitch or roll attitude.
The directional gyro indicator, which is often called the direction indicator or DI, provides a stable directional reference in azimuth
for maintaining accurate headings and for executing precise turns.
There is no magnetic element in the direction indicator, so it's not north seeking, and it must initially be synchronised
with the magnetic compass.
The synchronisation must be checked at regular intervals because of real and apparent gyro wander, sometimes called gyro drift.
The directional gyro indicator doesn't therefore replace the compass. Its stable, deadbeat indications are complementary to the north seeking capability
of the compass.
Having no magnetic element, the directional gyro indicator doesn't suffer from the compass turning and acceleration errors produced by the
vertical component of the Earth's magnetic field.
The directional gyro indicator employs a tied gyro, that's to say a gyro having freedom of movement in three planes
mutually at right angles to each other, but with the rotor axis maintained in the yawing plane of the aircraft.
This means that the rotor axis is horizontal in level flight, and because of gyroscopic rigidity, it provides the datum
from which heading can be measured.
The rotor is mounted in the inner gimbal, on bearings mounted in the outer gimbal, which has restricted freedom to
turn.
The outer gimbal can rotate through 360 degrees about the aircraft's vertical axis, on bearings within the case of the
instrument.
Note that the rotor axis, the inner gimbal axis, and the outer gimbal axis are mutually at right angles.
During a turn, the aircraft and the instrument case turn on the vertical axis bearings of the outer gimbal.
While the gyro rotor, gimbals and cylindrical indicating scale all remain fixed in azimuth because of gyroscopic rigidity.
Heading is indicated on the cylindrical indicating scale by a lubber line painted on a glass window in the instrument
case.
Another type, shown here, has a circular vertical card indicating scale geared to the outer gimbal, instead of the cylindrical
scale.
With earlier designs of directional gyro indicator, the rotor is driven by twin jets of air applied from the outer
gimbal ring as shown in here.
Suction is applied to the case of the instrument, and replacement air enters the case through a filter and is
ducted to the jets on the outer gimbal, which act on buckets cut in the rotor.
The jets not only spin the rotor, but also serve to maintain or tie the rotor axis in the yawing
plane of the aircraft.
If the axis of a gyro rotor departs from its chosen direction, it's said to wander.
If the axis of the gyro rotor wanders in the horizontal plane, as indicated by the horizontal arrow in this
diagram, it's said to drift.
A weight hung on the gimbal at point A will produce drift.
If, on the other hand, the rotor axis wanders in the vertical plane, as shown by this arrow, it's said
to topple.
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Manufacturing imperfections on the gyroscope cause small rates of random precession.
Other terms given to this precession are balance wander or, if the precession is in the horizontal plane, mechanical drift.
The imperfections concerned are uneven rotor bearing friction, unbalanced gimbals, and friction in the gimbal bearings.
In-flight turbulence may increase the effect of these imperfections.
The precession caused by application of an external force to a gyro is another example of real wander.
Consider a space, or free gyro, as shown here, with the rotor axis horizontal and aligned with the geographic meridian.
The rotor axis is indicating the direction of true north on the Earth and is also aligned with a point
at an infinite distance in space.
Some time later, when the Earth has rotated in space, the gyro rotor axis is still aligned with the same
fixed points in space, assuming no other disturbing forces, but no longer indicates the direction of north on the Earth.
It has therefore changed its alignment, according to the Earth-bound observer, by 70 degrees in this case at this
latitude.
This is apparent wander, or apparent drift, due only to the rotation of the Earth about its axis.
For thousands of years the oxide of iron called magnetite has been observed to attract small pieces of iron. This
property is known as magnetism.
Another property for which magnetite was known was its north-seeking capability.
If it was mounted on a piece of wood to allow it to float on water it would swing around
and align itself in a roughly north-south direction, so acting as a primitive compass.
In more recent history it was found that some metallic elements and alloys, mainly ferrous, iron and steel, could be
given these properties, bars of such magnetized material being known as magnets.
The field of a magnet is the space around it in which its magnetic influence is felt.
This may be illustrated by placing a piece of card over a bar magnet and scattering iron filings on it.
When the card is shaken or tapped the filings will take up the field pattern as shown in this diagram.
This same diagram illustrates that the lines of force traced by the iron filings converge towards small areas near the
ends of the magnet.
These two areas are called the poles of the magnet, and this is where the properties of magnetism are most
strongly displayed.
Magnets can be made in various shapes, but each magnet always has two poles.
A unit pole cannot exist. If a magnet is cut into two pieces, each piece will have two poles.
A freely suspended bar magnet, or a compass needle, in the Earth's magnetic field will align itself roughly north-south.
The end which points north is known as a north-seeking or red pole. The other end is a south
-seeking or blue pole.
The Earth behaves as though a huge permanent magnet was situated near its centre, producing a magnetic field over the
surface of the Earth.
You may remember we said that a freely suspended magnet in the Earth's magnetic field would align itself roughly north
-south.
For it to do this, its north and south-seeking ends would have to be attracted to poles of opposite
polarity.
Thus we portray this huge hypothetical permanent magnet inside the Earth with its blue pole under the Earth's magnetic north
pole and its red pole under the Earth's magnetic south pole.
This diagram shows that the poles of this hypothetical Earth magnet do not lie on the Earth's spin axis, this
lack of symmetry giving rise to magnetic variation.
The Earth's magnetic north pole, the blue pole, lies at present beneath northern Canada in the area around 70°N 95°W.
The magnetic south pole, the red pole, being below Antarctica at about 72°S.
A
compass is an instrument designed to indicate direction on the surface of the Earth.
relative to some known datum.
The magnetic compass uses the horizontal component of the Earth's field as its directional datum.
Unfortunately the Earth's field is normally not aligned with a true meridian, the most desirable datum from which to measure
direction.
The angular difference between true and magnetic meridians is called the magnetic variation.
The purpose of a magnetic steering compass in an aircraft is to indicate heading, the direction in which the aircraft
is pointing.
Magnetic influences, iron or steel components, electric currents, distort the Earth's field so that the compass magnetic assembly deviates from
the magnetic meridian.
This is called compass deviation.
The rules for applying variation and deviation to the compass heading indication in order to determine true heading are discussed
later in this lesson.
This part of this lesson deals with the direct indicating or direct reading magnetic compass, where the pilot directly reads
his heading in relation to a pivoted magnet assembly.
There are two basic types of direct reading magnetic compasses used in aircraft, the vertical card and less commonly the
grid ring compass.
The vertical card compass, which is also known as the B-type or E-type, is a direct reading compass
in general use.
It's usually the main magnetic heading reference in light aircraft and the standby compass in larger aircraft.
It consists of a circular compass card attached directly to the magnet assembly.
This combined unit is suspended in liquid within the compass bowl.
A vertical lubber line on the glass window of the bowl enables the heading to be read off the compass
card.
The P-type compass or grid ring compass is found on older aircraft. It's more accurate than the vertical card
compass and is more stable.
It is however heavier, bulkier and more expensive.
In addition, it can only be read in straight and level flight as the grid ring has to be unclamped
and aligned with the north reference before a reading can be taken against the lubber line.
The grid ring compass also differs from the vertical card compass in that it achieves a greater periodicity by the
addition of damping wires which also rotate through the compass liquid.
The Earth behaves as though a huge permanent magnet was situated near its centre, producing a magnetic field over its
surface.
The direction of that field at any given point can be indicated by a freely suspended magnet.
Such a magnet will align itself roughly in a north-south direction with its red pole towards the north magnetic
pole.
However, the poles of the Earth's magnetic field do move and are seldom coincident with the Earth's spin axis.
The position of the north magnetic pole is moving at almost 10 nautical miles per year and since it was
first located in 1831 it's moved approximately 625 miles and is presently close to Elif Ringnes island in the Canadian
Arctic islands which are 650 miles west of the geographic north pole.
That would make the magnetic lines of force lie like this.
And the freely suspended magnet will align itself like this.
The angle, measured in the horizontal plane between the magnetic meridian at a point and the true meridian at the
same point, is known as the magnetic variation.
Variation is designated west or east, depending on whether the magnetic pole lies to the west or to the east
of true north.
Variation can have any value from 0 to 180 degrees, the latter occurring on the true meridian, linking the north
geographical pole with the north magnetic pole.
A simple rhyme can be remembered which will help enormously in working out which way to apply variation to your
heading.
Variation west, magnetic best. Variation east, magnetic least.
Here an aircraft is flying due east by reference to points on the ground.
So if the variation is west of true north, say 11 degrees west, then add it to your aircraft's true
heading to get your correct magnetic heading of 101 degrees.
Conversely, if the variation is east of true north, say 11 degrees east, then subtract it from your aircraft's true
heading to get your correct magnetic heading of 079 degrees.
Remember, variation west, magnetic best. Variation east, magnetic least.
Except near the magnetic equator, where the lines of force are parallel to the surface, one end of the freely
suspended magnet will dip below the horizontal, pointing to the nearer pole.
To the south of the magnetic equator, the magnet's blue pole will be lower, whereas to the north, the red
pole will be lower.
The angle, measured in the vertical plane between the axis of the magnet and the horizontal is called the angle
of dip.
If the freely suspended magnet is moved either north or south of the magnetic equator, the dip gradually increases, reaching
about 66 degrees in the United Kingdom.
Over the Earth's magnetic poles, the dip is 90 degrees and the magnet is then vertical.
During a turn, the liquid in contact with the inside of the compass bowl tends to be dragged around with
the bowl.
Therefore, the liquid tends to swirl and rotate the magnet assembly with it, in the same direction as the aircraft's
turn.
Direct reading compasses are subject to large errors during linear acceleration or deceleration, or during a turn.
Turning and acceleration errors only occur where there is significant magnetic dip, so that except for a small liquid swirl
effect in turns, the errors are non-existent near the magnetic equator.
This diagram shows both the top view and the pilot's eye view of a pendulously suspended magnet, with residual dip,
in the northern hemisphere.
Note that while the aircraft is in cruising flight, the compass reading is correct.
The magnet is attached to the aircraft at the pivot point.
However, while the pivot is being accelerated, the magnet will try to stay where it is because of its inertia,
which acts at the magnet's centre of gravity.
The result will be that, even though the aircraft is flying in a straight line, the magnet will rotate, in
this case anti-clockwise, and the incorrect heading will be shown.
Notice that once the aircraft is in cruising flight, and the acceleration has finished, the magnet moves back to its
original position, and once more shows the correct reading.
Turning errors are maximum when turning through north and south, and ignoring liquid swirl, zero when turning through east and
west. The error increases with increase in magnetic latitude.
At the magnetic equator, the only turning error is due to liquid swirl.
Whenever the pilot turns through the nearer pole, that is the north pole in the northern hemisphere, or the south
pole in the southern hemisphere, the aircraft and compass rotate in the same direction, and the compass will be sluggish.
The pilot should undershoot the turn, that is, he should roll out early.
Whenever the pilot turns through the further pole, that is the south pole in the northern hemisphere, or the north
pole in the southern hemisphere, the aircraft and the compass rotate in opposite directions, so the compass will be lively.
The pilot should overshoot the turn, that is, he should roll out late.
Select any of these buttons to see how turning errors affect the compass.
Deviation is produced by the iron, steel and electrical components of the aircraft. It's the angle between the local magnetic
meridian and the direction in which the compass magnets are lying.
Deviation is named easterly, or plus, if the north-seeking red end of the magnet points to the east of
magnetic north.
If the north-seeking end points to the west of magnetic north, deviation is said to be westerly, or minus.
Deviation varies with heading, so it has to be measured on a series of different headings. This is usually done
by conducting a compass swing.
Once deviation has been reduced as far as possible, the residual deviation is recorded on a compass deviation card, which
is located in the aircraft.
During the swing, normal flying conditions should be simulated as far as possible, with engines running, electrical and radio services
switched on, and the aircraft in a level flight attitude.
It's obviously most important that no ferromagnetic objects such as tools or watches should be placed near the compass, as
this would introduce unknown amounts of deviation.
Here we see that we are flying with a compass reading of 270 degrees. The compass reading is the magnetic
heading plus or minus the deviation on that particular heading.
The deviation card makes it easy for us to fly an accurate magnetic heading. In this case, it's telling us
that to fly on a magnetic heading of 270 degrees, we must steer 269 degrees.
The basic method of determining deviation is to compare the aircraft's heading compass reading with magnetic heading, as defined by
high quality land or datum compass.
This comparison of aircraft compass and magnetic datum readings is carried out in an area selected specifically for this purpose.
Major compass errors are corrected by adjusting small magnets mounted in the compass system.
Any residual errors that remain are shown on a compass deviation card displayed in close proximity to the compass in
the cockpit.
There are several occasions when an aircraft might require compass swing. Among these are
When compass components are installed or replaced
Whenever the accuracy of the compass is in doubt
After a maintenance inspection if required by the schedule
After a significant aircraft modification, repair or replacement involving magnetic material
If the aircraft has been struck by lightning
Check that there is no obvious damage to the compass body such as dents or cracks
Any lighting system should be checked, as should the efficacy of luminous paint.
A compass liquid should be checked and be free from sediment and discolouration
Either of which would indicate corrosion, which would result in increased pivot friction
A compass liquid should also be free from bubbles, which would probably indicate a leaking seal
Turbulence and manoeuvres would cause any bubbles to move about, creating eddies which could disturb the magnet system
The compass reading can also be checked for gross errors when you first enter the aircraft
and make sure that the compass deviation card is placarded conveniently close to the compass
Check the compass reading while taxiing the aircraft
Remember, the compass reading should decrease when you're turning left and increase when you're turning right
Just before take-off, check the compass reading against the runway QDM
It would be very difficult to fly a modern light aircraft safely without its engine instruments
Some engine instruments are what are termed engine condition indicators
These include the oil pressure and oil temperature gauges and the exhaust gas temperature gauge
They literally show the actual condition of the engine
Other engine instruments are called performance indicators
Within this group are placed the engine RPM gauge and the manifold pressure gauge
They are intended to show how well the engine is performing
Furthermore, without engine instruments, effective management of the engine would be impossible
The engine would not give long and efficient service and faulty operation would be commonplace
It's thus essential that engine instruments be fitted to aircraft and that they function correctly
The pilot must also be assiduous in monitoring the engine instruments
Piston engines are heat engines
The power they produce is directly proportional to the heat released during the combustion of the fuel
Engine components and systems are designed to withstand certain temperatures
If the temperature limits are exceeded, the components may fail
To allow safe operation, the engine temperatures must be monitored
The following temperatures are monitored in piston engines
Oil, exhaust gas and cylinder head
Monitoring the engine oil temperature is important from at least two points of view
Firstly, the oil itself will only retain its lubricating properties up to a certain maximum temperature
Above that temperature, the oil breaks down and the moving parts of the engine begin rubbing against each other and
getting even hotter
Engine failure quickly follows
Secondly, the engine expends a percentage of its power forcing the oil into the bearings
The colder the oil, the larger that percentage of what is essentially wasted power will be
If the aircraft is to perform safely, the engine oil temperature must reach a certain level
before the pilot can be sure that his engine is delivering the power to the propeller that the manufacturers intended
Oil temperature is normally sensed where the oil enters the engine, after it has exited the oil cooler
In a typical engine oil temperature indicating system, the indicator is powered by the aircraft electrical system
An electrical resistant type thermo bulb installed in the engine oil pump housing measures the temperature of the oil entering
that unit
The temperature reading is transmitted to the indicator, usually in degrees Celsius
An exhaust gas temperature gauge can show you whether your engine combustion chamber is running too hot or too cold,
too lean or too rich
An exhaust gas temperature probe is installed about 4 inches from the cylinder head on the exhaust system
The probe is the junction of two dissimilar metals. This is called a thermocouple
When a thermocouple is heated, a voltage is produced which is proportional to the temperature at the junction
The gauge is a millivolt meter which is calibrated in degrees Celsius or Fahrenheit
Although the exhaust gas temperature gauge can help in troubleshooting a problem on the engine, it's primarily a fuel management
instrument
The cylinder head temperature gauge is an engine instrument designed to protect the engine against its enemy, excessive heat
The principle of operation of the cylinder head temperature gauge system is similar to that of the exhaust gas temperature
gauge
The probe is mounted on the engine cylinder head and this sends a voltage to the gauge in the cockpit
which is proportional to the temperature
Most general aviation engines sample the cylinder head temperature of the hottest cylinder
The hottest cylinder is determined by extensive flight tests carried out by the engine manufacturers
The function of many aircraft engine systems relies on liquids and gases whose pressure must be measured and indicated
The gauges and indicating systems fall into two categories, direct reading and remote indicating
Direct reading describes the type of instrument used when the fluid that is being sampled is fed directly to the
interior of the instrument positioned in the cockpit
Remote indicating is where a separate sensing element is connected to a pressure source at some remote point
and the information required is transmitted electrically to the instrument in the cockpit
Pressure is defined as force per unit area
In engine instrument terms it's normally indicated in either pounds per square inch or inches of mercury
In connection with engine pressure measurement we are concerned with the following terms
Absolute pressure and gauge pressure
Most pressure gauges measure the difference between the absolute pressure and the atmospheric pressure
To actually measure pressure in a system, elastic pressure sensing elements are used in which forces can be produced by
applied pressures and converted to mechanical movement
The movement can then operate a direct reading gauge or an electrical transmitter
The sensing elements commonly used are diaphragms
capsules
bellows
and Borden tubes
Diaphragms consist of corrugated circular metal discs which are secured at their edge and when pressure is applied they are
deflected as shown here
Diaphragms are used to measure relatively low pressures
Capsules are made up of two diaphragms placed together and joined at their edges
The device thus constructed can then be used to either form a sealed chamber which is called an aneroid capsule
or a chamber which is communicated to a pressure source which is called a pressure capsule
Capsules, as diaphragms, are also used to measure low pressure but capsules are more sensitive to small pressure changes than
are diaphragms
The bellows type element can be considered as an extension of the corrugated diaphragm principle
It may be used for high, low or differential pressure measurement
The Borden tube is about the oldest of the pressure sensing methods
The element is essentially a length of metal tube which has an elliptical cross section
The tube is shaped to represent the letter C
One end of the tube is sealed. This is called the free end
The other end is connected to the pressure source and fixed so that it cannot move in relation to the
instrument base
When pressure is applied to the tube it tries to straighten
This movement is magnified by levers and used to drive an indicator pointer
The Borden tube can be manufactured to indicate high or low pressure but it is normally associated with higher pressures
such as engine oil pressure
Oil pressure is sensed at the outlet of the oil pressure pump
Normally the oil pressure of an engine is maintained constant by an oil pressure relief valve
There are however situations which may affect the engine or the oil to the extent that it is beyond the
ability of the pressure relief valve to maintain the oil pressure within the prescribed limits
It's at these times that the oil pressure gauge, if it's used in conjunction with the oil temperature gauge, can
indicate to the pilot the best course of action to retrieve the situation
The manifold absolute pressure gauge of a piston engine uses two bellows to measure both ambient atmospheric pressure and the
pressure in the inlet manifold
Note that this gauge measures absolute pressure and is calibrated in inches of mercury
When the engine is running this gauge may indicate less than atmospheric pressure
Earlier versions of this gauge were calibrated to read boost in pounds per square inch and called boost pressure gauges
When the engine is not running both types of gauge, the manifold absolute pressure gauge and the boost gauge, will
read ambient atmospheric pressure
While the aircraft is stationary on the ground this particular indication is called static boost
A fuel pressure gauge may be fitted to fuel injection fuel systems and can be used to indicate metered fuel
pressure
In some cases, by calibrating it in a suitable manner, the fuel pressure gauge can enable the pilot to adjust
the air-fuel mixture to suit the aircraft's altitude and the engine power setting
The measurement of the quantity of fuel in the tanks of an aircraft fuel system is an essential requirement
The simplest form of volume indication is a float system
Early aircraft had a float which sat on top of the surface of the fuel
Attached to the float was a piece of wire that protruded out of the top of the fuel tank
As the fuel level reduced, so the wire disappeared from view
There have been many variations of this system
On light aircraft, the most common of these uses a float which moves to position a wiper on a variable
resistor
This in turn alters the current flowing to an indicator, moving a pointer over a scale calibrated in volume
The disadvantage of this system is that the indication is not linear and there is no provision for making adjustments
for system accuracy
The gauge is set to be accurate at the low and empty positions
The system is also subject to errors whenever the aircraft manoeuvres and the attitude changes
As well as the quantity of fuel measured, the rate of fuel flow can be shown
The fuel flow meter can display volume flow or mass flow
A simple flow meter can be an adaptation of a pressure gauge
This system is used on many light aircraft piston engine fuel injection systems
The measurement of engine speed is of vital importance, since together with other parameters, accurate control and monitoring of the
engine can be achieved
On piston engines, it's crankshaft speed that is measured
The RPM indicator is called a tachometer or tacho for short
The basic method of measuring engine rotational speed on piston engines is the mechanical magnetic tachometer
The mechanical tachometer consists of a flexible driveshaft and the tacho indicator
One end of the flexible driveshaft is connected to the tacho indicator in the cockpit
The other end goes to the accessory drive casing on the engine, where it's driven through gears from the crankshaft
The input drive causes a magnet to rotate
The magnet rotates inside a copper or aluminium drag cup
This induces eddy currents in the aluminium drag cup
Eddy currents are caused by the magnetic field of the moving magnet, acting on the electrons of any metal in
the vicinity of the magnet
The eddy currents themselves generate a magnetic field, which interacts with the magnetic field of the magnet
This interaction causes a torque, or turning moment, which is established to turn the drag cup in the same direction
as the permanent magnet
A shaft extends from the drag cup and is connected to a pointer
The turning motion of the pointer is against the tension of a hairspring, which controls the drag cup position and
hence the position of the pointer
The flexible drive is driven at reduced speed, but true speed will be shown on the indicator
The indicator incorporates compensation devices for changes in temperature
Left click on the blue slider button situated below the tacho generator diagram with your mouse, and move it right
and left to see how the device works
Notice how the spring restricts the drag cup as it turns the pointer
The United Kingdom Air Navigation Order stipulates that all United Kingdom air navigations have the right to fly
The United Kingdom registered aircraft should meet specific airworthiness requirements
The most pertinent of these documents are the following
The Certificate of Airworthiness
The Certificate of Registration
The Flight Manual
and the Maintenance Documents
The Certificate of Type Approval
When an aircraft manufacturer wishes to put a new type or model of aircraft on the market, he must first
obtain a Certificate of Type Approval
The prototypes of the particular aircraft type must pass many stringent tests before this can occur
But when it does happen, the Certificate of Type Approval is issued by the Aviation Authority in the country of
manufacture
The type certificate is retained by the aircraft manufacturer
Tests for the issue of a Certificate of Type Approval cover safety, engineering, reliability and aircraft performance
The United Kingdom Civil Aviation Authority does not normally carry out a UK type certification for foreign aircraft of conventional
design that weigh less than 6,000 pounds or 2,730kg
but accepts the aircraft on the basis of the foreign type certificate
The United Kingdom Civil Aviation Authority carries out a validation exercise in accordance with the procedures detailed in British Civil
Airworthiness Requirements, BCARs
The level of CAA investigation of the aircraft depends on the intended operation and certification category, and the associated equipment
fit and performance information
The United Kingdom Civil Aviation Authority requires that all United Kingdom owned or operated aircraft be registered with them
As soon as this is done, the United Kingdom Civil Aviation Authority issues a Certificate of Registration to the owner
and the aircraft is assigned its registration letters, by which it may be recognised
In the United Kingdom, four registration letters follow the United Kingdom nationality marking G, like these shown here on Golf
Bravo X-Ray Yankee Tango
The Certificate of Airworthiness, usually referred to in its shortened form as the C of A, is issued to a
particular aircraft by the Civil Aviation Authority for a specific period
normally three years for an aircraft of less than 6,000 pounds, 2,730kg
The Certificate of Airworthiness confirms that the aircraft is certified to an internationally acceptable standard in terms of safety, design,
construction and performance, etc.
A Certificate of Airworthiness is issued to an aircraft to allow it to operate within a specific category, as long
as it complies with the appropriate airworthiness requirements
Unless the aircraft has a permit to fly, which we will mention later, or is undergoing a test flight, it
cannot fly without a valid Certificate of Airworthiness
The Certificate of Airworthiness states any restrictions and conditions of issue. It also states that both the Flight Manual and
the Pilot's Operating Handbook are part of the Certificate of Airworthiness
Ignoring any of the restrictions or conditions laid down in the Certificate of Airworthiness, or the procedures and limitations in
the Flight Manual, or exceeding the period of validity, will automatically invalidate the Certificate of Airworthiness
The categories and purposes that are apposite to us are
Transport, Passengers, Any Purpose
Transport, Cargo, Any Purpose except the Public Transport of Passengers
Aerial Work, Any Purpose except Public Transport
Private, Any Purpose except Public Transport or Aerial Work
Special, Any Purpose other than Public Transport specified on the C of A, but usually precluding the carriage of passengers
Aircraft are further categorised according to the manoeuvres that they are allowed to perform. These are
Normal Category
Below 5,700kg and non-aerobatic, manoeuvres limited to stalls and steep turns of 60 degrees
Typical limit load factors plus 2.5g and minus 1g
Utility Category
As for Normal Category plus limited aerobatics
Typical limit load factors are plus 4.5g and minus 1.8g
Acrobatic Category
Fully aerobatic
Typical limit load factors are plus 6g and minus 3g
Flight Manual
The Flight Manual forms part of the Certificate of Airworthiness for a particular aircraft
The United Kingdom Civil Aviation Authority certifies an aircraft for use in its own area of responsibility
and may add amendments to the Flight Manual
These amendments are called the Flight Manual Supplement
and the pilot must comply with all the requirements, procedures and limitations set out in the Flight Manual
as amended by the Flight Manual Supplement with respect to the operation of the aircraft
A United Kingdom Civil Aviation Authority Flight Manual Supplement overrules the information in the original Flight Manual or Pilot's Operating
Handbook
The Flight Manual and the United Kingdom Civil Aviation Authority Flight Manual Supplement must be carried in the aircraft at
all times
In order that the pilot can have some other reference to the data applicable to his aircraft
the Pilot's Operating Handbook has been derived from the Flight Manual
The same procedures, techniques, limitations and performance standards which are contained in the Flight Manual
are to be found in the Pilot's Operating Handbook which can be the pilot's personal property
As an alternative to a Certificate of Airworthiness certain aircraft in the United Kingdom may be granted a permit to
fly
It's a national permission to fly issued by the United Kingdom Civil Aviation Authority for use within the United Kingdom
only
and is in recognition of the fact that amateur-built aircraft and some vintage types which do not qualify for
a C of A would otherwise be grounded
It may also be issued to allow an aircraft a position in order to undergo a test or inspection
In the United Kingdom the inspection and maintenance requirements of aircraft which have a permit to fly are in the
main managed by the Popular Flying Association, the PFA
A permit to fly restricts the operation of the aircraft as follows
Firstly the aircraft shall be used for recreational purposes only, that is not for public transport or for aerial work
such as aerial photography or banner towing
The one exception however is that PFA aircraft may be flown for public exhibition or demonstration such as at air
shows, even when this does constitute aerial work
Secondly, except to the extent necessary in order to take off and land in accordance with normal aviation practice, PFA
aircraft are not allowed to fly over any assembly of persons or congested area at any height
Thirdly, PFA aircraft are permitted to fly in daytime and under visual flight rule flight conditions only, that is not
at night and not in accordance with instrument flight rules
As stated earlier, a permit to fly permits flight in the United Kingdom only, it is not an internationally recognised
certificate of airworthiness and does not infer a right of flight in the airspace of any other country
However, every permit to fly carries an exemption that allows the aircraft to leave UK airspace, which the Air Navigation
Order would otherwise prohibit
Note however that this exemption still does not represent a permission to enter the airspace of another country
All aircraft in the transport, aerial work and private categories are to be maintained in accordance with either an approved
maintenance schedule or, if its maximum take-off weight authorised does not exceed 6,000 pounds, 2,730kg, the light
aircraft maintenance schedule
These both require that a system of regular checks and inspections is carried out by licensed and approved persons
Furthermore, logbooks must be kept in respect of the engine and the airframe and, if the aircraft is fitted with
a constant speed propeller, it too must have a logbook
A typical maintenance schedule will include
A daily inspection, sometimes called the Check A. This is a thorough pre-flight inspection of the aircraft which can
be carried out by the pilot
A 50 hour check, which is a detailed inspection of the aircraft
A 150 hour check, which is much more detailed than the 50 hour check
The Annual Inspection, which is a major inspection of the engine, the airframe and all of the aircraft systems and
components
The Star Annual Inspection, which is an expanded annual inspection carried out every three years. This inspection includes a test
flight to ensure that the aircraft is performing to the standards set out in its flight manual
A valid Certificate of Maintenance review must be enforced for an aircraft with a Transport Category Certificate of Airworthiness
This confirms that the aircraft has been maintained in accordance with its maintenance schedule and also that any airworthiness directives
released within the previous 12 month period have been complied with
This also applies to any manufacturer's service bulletins which may have been released during the same period
Airworthiness directives may often be used to make a manufacturer's service bulletin compulsory
Following any maintenance carried out on an aircraft that has a Certificate of Airworthiness, a Certificate of Release to Service
is issued by a licensed engineer
This certifies that the work carried out on the aircraft has been done in accordance with both the Civil Aviation
Authority and the manufacturer's procedures
Work carried out by a pilot or the owner of the aircraft does not require a Certificate of Release to
Service
The maintenance which can be carried out by the pilot or owner is described in some detail later in this
lesson
An aircraft with a Transport Category Certificate of Airworthiness must have a technical log
The technical log is used by the pilot to record take-off and landing times and it must also be
signed by him after each flight
Any defects apparent after the completion of a flight can be entered in the technical log and the maintenance staff
will use it to record any work carried out to rectify those defects
Although in general all maintenance on an aircraft with the United Kingdom Certificate of Airworthiness must be carried out and
inspected by a licensed aircraft engineer
The United Kingdom Civil Aviation Authority has made provision for certain minor repairs and services to be performed by owners
or operators who are not licensed engineers
This applies if the aircraft is less than 6,000 pounds, 2,370 kilograms, and is in a category other
than public transport
A list of those tasks allowed to be performed by a private pilot's license holder who is not a licensed
engineer is published in the Air Navigation General Regulations No. 16
That list is reproduced here
Replacement of landing gear tires, landing skids, or skid shoes
Replacement of elastic shock absorber cord units on landing gear where special tools are not required
Replacement of defective safety wiring or split pins, excluding those in engine, transmission, flight control, and rotor systems
Patch repairs to fabric not requiring rib stitching or the removal of structural parts or control surfaces if the repairs
do not cover up structural damage and do not include repairs to rotor blades
Repairs to upholstery and decorative furnishing of the cabin or cockpit interior when repair does not require dismantling of any
structure or operating system or interfere with an operating system or affect the structure of the aircraft
Repairs not requiring welding, deferrings, non-structural cover plates, and cowlings
Replacement of side windows where that work does not interfere with the structure or with any operating system
Replacement of safety belts or safety harnesses
Replacement of seats or seat parts not involving dismantling of any structure or of any operating system
Replacement of bulbs, reflectors, glasses, lenses, or lights
Replacement of any cowling not requiring removal of the propeller, rotors, or disconnection of engine or flight controls
Replacement of unserviceable sparking plugs
Replacement of batteries
Replacement of wings and tail surfaces and controls, the attachments of which are designed to provide for assembly immediately before
each flight and dismantling after each flight
Replacement of main rotor blades that are designed for removal where special tools are not required
Replacement of generator and fan belts designed for removal where special tools are not required
Replacement of VHF communication equipment, being equipment which is not combined with navigation equipment
In addition to this list, the holder of a private pilot's licence may also carry out the 50 hour inspection
of an aircraft in the private category if he or she is the owner or operator of the aircraft
Details of any work carried out under any of the circumstances just mentioned must be entered in the aircraft logbook
and signed for by the person carrying out the work, either an engineer or a pilot qualified on type
If any adjustments are made to the flight controls or the engine controls, these adjustments would normally be required to
be checked by two licensed personnel, either engineers or inspectors, to confirm that the work has been properly carried out
and that the controls are functioning properly
This is termed a duplicate control system inspection
However, British Civil Airworthiness requirements state that should a minor adjustment of a control system be necessary when the aircraft
is away from base and where the initial part of the inspection has been done by a licensed engineer, the
second part of the inspection can be carried out and certified by a pilot licensed for the type of aircraft
concerned
The daily inspection, also known as the Check A in the United Kingdom, must be carried out prior to the
first flight of the day, although in practice it's carried out before each flight. This check can be carried out
by the pilot
The pilot must confirm that the aircraft he or she is about to fly has valid insurance cover
Although a Certificate of Insurance is not a part of the Certificate of Airworthiness, it's only common sense to make
sure that you and your aircraft are covered for any eventuality